INERTIAL INSTRUMENT SYSTEM 

FOR AERIAL SURVEYING

U.S. GEOLOGICAL SURVEY PROFESSIONAL PAPER 1390
Inertial Instrument System for 

Aerial Surveying

By RUSSELL H. BROWN, with contributions by WILLIAM H. CHAPMAN, 

WILLIAM F. HANNA, CHARLES E. MONGAN, U.S. Geological Survey, and 

JOHN W. HURSH, The Charles Stark Draper Laboratory, Inc.

U.S. GEOLOGICAL SURVEY PROFESSIONAL PAPER 1390

UNITED STATES GOVERNMENT PRINTING OFFICE, WASHINGTON: 1987
DEPARTMENT OF THE INTERIOR 

DONALD PAUL HODEL, Secretary

U.S. GEOLOGICAL SURVEY 

Dallas L. Peck, Director

Library of Congress Cataloging-in-Publication Data

Brown, Russell H.

Inertial instrument system for aerial surveying.

(U.S. Geological Survey professional paper ; 1390)

Bibliography : p. 95

Includes index.

Supt. of Docs, no.: I 19.16:1390

1. Aerial photogrammetry Equipment and supplies. 2. Inertial navigation systems I. Title. II. Series.

TA593.35.B76 1987 526.9'823 85-600134

For sale by the Books and Open-File Reports Section, U.S. Geological Survey, 

Federal Center, Box 25425, Denver, CO 80225
FOREWORD

In the late 1960's and early 1970's, the much-heralded and precise performance of inertial guidance systems in space exploration challenged earth scientists to seek ways of turning those systems to their advantage in the nearearth environment. Russell Brown and Charles Mongan sensed and responded 

to that challenge. Over a 2-year period (1972-74), they shaped a workable 

research task and secured the endorsement of the top U.S. Geological Survey 

managers. As a result, the Geological Survey was committed to the development of a pioneering inertial instrument system for aerial surveying, through 

contract arrangements with The Charles Stark Draper Laboratory.

The formal contract work began in mid-1974, at a time when a few military 

and civilian agencies were successfully operating less-precise generations of 

inertial guidance systems in motor vehicles and helicopters to establish control 

lines for topographic surveys. The Geological Survey's research commitment 

was to move rapidly beyond those techniques and to field a new and more 

precise inertial guidance system capable of obtaining similar or better control 

lines from a fixed-wing aircraft. The potential savings in dollars, manpower, 

and time in using the new system for field studies in the Geological Survey 

Operating Divisions more than justified this research commitment.

After nearly a decade of work, a prototype model of this new instrument 

system has been successfully flight tested. Its demonstrated performance has 

exceeded in every respect the stringent design specifications. Although the full 

range in performance capabilities remains to be explored, the system will 

provide highly efficient, cost-effective support for scientific programs of the 

Geological Survey, including production of topographic maps, investigations of 

water resources, and assessment of mineral resources and geologic hazards. 

The system also has immediate potential to support the research and operational missions of other Federal, State, local, and private organizations.

The Geological Survey is pleased to have played a major developmental role 

in this important scientific achievement.

Dallas L. Peck 

Director

PREFACE

The purpose of this report is to describe, for the nonspecialist earth scientist 

or engineer, an inertial guidance or navigation system that will enable use of 

relatively light aircraft for efficient nationwide data-gathering in geology, 

hydrology, terrain mapping, and gravity-field mapping. Development of this 

airborne instrument system is a logical consequence of the phenomenal postWorld-War-II growth in the science and technology of inertial guidance systems. That growth, although obviously predestined by the genius of men like 

Dr. Charles Stark Draper, was accelerated by the onset of the international 

space race in the 1960's.

By the early 1970's, new generations of gyroscopes and accelerometers key 

components in inertial systems had been designed, built, laboratory proven, 

and field tested sufficiently to allow application of inertial systems to certain 

high-precision field measurement processes in the earth sciences. Prime examples of successful applications in that period include the military development 

of a position and azimuth determining system, to run survey control lines, and 

the development for the National Aeronautics and Space Adminstration of the 

inertial navigators that guided the round-trip Apollo flights to the Moon. In 

these applications, however, the precision levels were lower than what the U.S. 

Geological Survey would require for its use.

The first inertially guided transcontinental flight, from Boston to Los Angeles, occurred in February 1953 with an Air Force B-29 carrying an inertial 

navigator that weighed more than a ton (over 900 kilograms). Dr. Draper 

(system designer), John Hursh (system operator), Charles Collins (pilot), and 

Irving Levin (flight engineer) were among The Charles Stark Draper Laboratory contingent that comprised that historic crew, and all have been involved 

actively in the current Geological Survey project. The present-generation inertial navigators in regular airline use individually weigh less than 100 pounds 

(45 kilograms) and can be accommodated in a space roughly equivalent to a 

desk file drawer.

In a timespan of about two decades, this kind of evolution in inertial system 

design, with equally graphic improvements in precision, has made the current 

instrument-development project technically feasible. Not only did feasibility 

crystallize, but the bright prospects that relate to the continuing evolutionary 

trend in inertial navigator design virtually guarantee that an instrument 

system built to satisfy the present Geological Survey specifications for precision can be upgraded periodically to meet even more stringent specifications 

whenever cost-effectiveness criteria dictate.

The instrument system described in this volume capitalizes not only on 

virtual state-of-the-art inertial guidance technology but also on similarly advanced technology for measuring distance with electromagnetic radiating 

devices. The distance measurement can be made with a transceiver beamed at 

either a cooperative target, with a specially designed reflecting surface, or a 

noncooperative target, such as the Earth's surface. The instrument system 

described herein features components that use both techniques. Thus, a laser 

tracker device, which updates the inertial guidance unit or navigator in flight, 

makes distance measurements to a retroreflector target mounted at a groundcontrol point; a laser profiler device, beamed vertically downward, makes distance measurements to the Earth's surface along a path that roughly mirrors 

the aircraft flight path. In both technological domains, inertial guidance and 

electromagnetic distance measuring, the equipment components selected for 

use in this new instrument system were advanced well beyond levels that were 

then commercially available.

CONTENTS

	Page

Foreword ------------------------------------------------ in

Preface -------------------------------------------------- v

Commonly used inch-pound terms and their

preferred SI equivalents -------------------------------- ix

List of abbreviations -------------------------------------- ix

Abstract ------------------------------------------------- l

Introduction --------------------------------------------- 2

Developmental incentives ----------------------------- 2

Developmental strategy ------------------------------- 2

Developmental history -------------------------------- 3

Acknowledgments ---------------------------------------- 6

Instrument system --------------------------------------- 7

Overall system concept ------------------------------- 7

Inertial navigator ------------------------------------ 8

Inertial measurement unit -------------------------- 9

Gyroscopes ------------------------------------- 10

Accelerometers --------------------------------- 15

Gimbal system --------------------------------- 19

Resolvers -------------------------------------- 19

Feedback loops --------------------------------- 23

Electronics ------------------------------------- 24

Temperature control ---------------------------- 25

Performance dynamics ------------------------------ 27

External measurement requirement ----------------- 35

Laser tracker -------------------------------------- 36

Computer ------------------------------------------ 46

Physical attributes ----------------------------- 47

Subsystems and data interfaces ----------------- 50

Computer Continued

Peripheral devices -------------------

Software ----------------------------

Laser profiler -----------------------------

Color television camera --------------------

Special role of gravity -------------------------

Aircraft for deployment of the instrument system 

Operation of the instrument system ------------

Ground control ----------------------------

Flight planning ---------------------------

Flight mission ----------------------------

Maintenance and service -------------------

Postflight data processing ----------------------

Performance-evaluation flights -----------------

Overall objectives -------------------------

Calibration range -------------------------

Specified flights ---------------------------

Flight results -----------------------------

Flights 1 and 2 ------------------------

Flights 3 and 4------------------------

Uses for the instrument system ----------------

Hydrologic studies -------------------------

Map production tasks ----------------------

Geologic studies ---------------------------

Future improvements in the instrument system - 

Selected references ----------------------------

Symbols and dimensions -----------------------

Index -----------------------------------------

Page

53

54

66

73

74

76

77

79

81

83

84

84

85

85

85

87

89

89

89

91

91

92

93

93

95

96

98

ILLUSTRATIONS

FIGURE 1. Fiscal year history of work on U.S. Geological Survey contract by Draper Laboratory, Cambridge, Mass., 1975-81 --

2. Frequency response characteristics of inertial navigator and tracker to vehicle motion ---------------------------

3. Schematic of gyro and gimbal-support structure ---------------------------------------------------------------

4. Schematic of single-degree-of-freedom gyro --------------------------------------------------------------------

5. Photograph showing the single-degree-of-freedom floated gyro --------------------------------------------------

6. Schematic of gyro and accelerometer arrays for aerial profiling of terrain system --------------------------------

7. Schematics of acceleration-measuring devices ------------------------------------------------------------------

8. Schematic view of the single-axis pendulous integrating gyro accelerometer -------------------------------------

9. Photograph showing the pendulous integrating gyro accelerometer ----------------------------------------------

10. Schematic of support structure for inertial measurement unit ---------------------------------------------------

11. Resolver schematic and nomenclature -------------------------------------------------------------------------

12. Photograph showing disassembled multispeed resolver for gimbal-support structure ------------------------------

13. Schematized feedback loops ----------------------------------------------------------------------------------

14. Schematic diagram of electrical circuits for stabilization subsystem of inertial measurement unit -----------------

15. Thermal control schematic for inertial measurement unit ------------------------------------------------------

16. Support and thermal control structure for inertial measurement unit -------------------------------------------

17. Orientation of locally level three-coordinate reference scheme at selected points on an idealized ellipsoidal Earth --

18. Schematic of support structure for inertial measurement unit and tracker ---------------------------------------

19. Photograph showing the inertial measurement unit, laser tracker, gimbal-support structure, and housing ---------

20. Optical layout for laser tracker -------------------------------------------------------------------------------

21. Analysis of the range equation for the laser tracker ------------------------------------------------------------

22. Principal reference frequencies used in the aerial profiling of terrain system -------------------------------------

23. Cone-shaped air space within which the tracker can see a retroreflector target -----------------------------------

24. Photograph showing principal components and final assembly of the U.S. Geological Survey-designed retroreflector

25. Functional structure of the aerial profiling of terrain system computer with connections to subsystems and

	peripheral devices ---------------------------------------------------------------------------------------

26. Photograph showing the packaged flight computer for the aerial profiling of terrain system -----------------------

27. Photograph showing the operator's keyboard and plasma display terminal for the aerial profiling of terrain system

28. Photograph showing the line printer for the aerial profiling of terrain system ------------------------------------

29. Photograph showing the magnetic tape controller for the aerial profiling of terrain system -----------------------

30. Photograph showing the magnetic tape recorder for the aerial profiling of terrain system -------------------------

31. Hypothetical sequence of the aerial profiling of terrain system modes for a flight mission -------------------------

32. Structure graph of the master operating program for the aerial profiling of terrain system ------------------------

Page

5

8

10

11

13

15

16

17

18

20

21

22

23

25

26

28

30

37

39

40

42

44

45

46

47

49

52

53

54

55

56

61
INSTRUMENT SYSTEM FOR AERIAL SURVEYING

33. Recurrence intervals and callup sequence for the five synchronous tasks in the aerial profiling of terrain system ---- 66

34. Data flow for the STARTUP AND TEST mode of the software for the aerial profiling of terrain system ------------- 67

35. Data flow for the CALIBRATE AND ALINE and STANDBY modes of the software for the aerial profiling of

	terrain system -------------------------------------------------------------------------------------------- 68

36. Data flow for the four flight modes of the software for the aerial profiling of terrain system ------------------------ 69

37. Photograph showing basic chassis with optics and some electronics for the laser profiler --------------------------- 71

38. Analysis of the range equation for the laser profiler ------------------------------------------------------------- 72

39. Photograph showing color television camera for the aerial profiling of terrain system ----------------------------- 74

40. Photograph showing video recorder and control box for the aerial profiling of terrain system ----------------------- 75

41. Photograph showing time-date and time-code generators for the aerial profiling of terrain system ------------------ 76

42. Photograph showing DeHavilland Twin Otter aircraft ----------------------------------------------------------- 79

43. Arrangement of aerial profiling of terrain system components in DeHavilland Twin Otter aircraft ----------------- 80

44. Flight-path layout for profiling the Farmington River flood plain near Avon, Conn ------------------------------- 82

45. Information flow through postmission data-processing software -------------------------------------------------- 84

46. Control diagram and retroreflector sites for calibration range ---------------------------------------------------- 86

47. Flight plan for performance-evaluation flights 1 and 2 ---------------------------------------------------------- 88

48. Flight plan for performance-evaluation flight 3 ----------------------------------------------------------------- 88

49. Flight plan for performance-evaluation flight 4 ----------------------------------------------------------------- 88

TABLES

	Page

TABLE 1. Package size and power requirements for the aerial profiling of terrain system computer and peripherals __..---------- 48

2. Computer-action summary for the 14 special tasks in the aerial profiling of terrain system -------------------------- 62

3. Factors bearing upon aircraft selection ---------------------------------------------------------------------------- 77

4. Performance characteristics, DeHavilland Twin Otter aircraft ------------------------------------------------------- 78

5. Position coordinates and gravity values for retroreflector sites in the calibration range -------------------------------- 87

6. Results of performance-evaluation flights 1 and 2 ------------------------------------------------------------------ 90

7. Results of performance-evaluation flights 3 and 4 ------------------------------------------------------------------ 91
Commonly used Inch-Pound terms and their 

preferred SI equivalents

Multiply inch-pound units By To obtain SI equivalent

Length

inch (in.) 

foot (ft) 

yard (yd) 

mile (mi) 

nautical mile (nmi)

2.540 x 10*a 

3.048 x 10-! 

9.144 x 10-1 

1.609 

1.852*

millimeter (mm) 

meter (m)** 

meter (m) 

kilometer (km) 

kilometer (km)

Area

square foot (ft2 ) 9.290 x 10-2 square meter (m2)

Volume

cubic foot (ft3 ) 2.832 x 10-2 cubic meter (m3 )

Mass

ounce, avoirdupois (oz) 

pound, avoirdupois (Ib)

2.835 x 10 

4.536 x 10-!

gram (g) 

kilogram (kg)**

Force

ounce-force (ozf) 

pound force (Ibf)

2.780 x lO- 1 

4.448

newton (N) 

newton (N)

Pressure

pound per square inch (lb/in. 2 ) 

pound per square foot (lb/ft2 ) 

bar 

millimeter of mercury (0°C) 

inch of mercury (0°C)

6.895 x 103 

4.788 x 10 

1 x 105* 

1.333 x 102 

3.386 x 103

pascal (Pa) 

pascal (Pa) 

pascal (Pa) 

pascal (Pa) 

pascal (Pa)

Torque

ounce-inch (oz-in.) 

pound-foot (lb-ft) 

dyne centimeter (dyn-cm)

7.062 x lO-3 

1.356 

1 x 10-7*

newton meter (N-m) 

newton meter (N-m) 

newton meter (N-m)

Temperature

degree Fahrenheit (°F) Temp °C = (Temp °F - 32)(5/9) degree Celsius (°C)

Illumination

foot-candle (ft-c) 1.076 x 10 lux (Ix)

*Exact conversion factor. 

**Basic SI unit.

aExact conversion, except for geodetic surveying in the United States for which the conversion factor is 2.540005.

List Of Abbreviations

Abbreviation 

APT(S) 

DC 

DMA 

GaAs 

gyro 

Hz 

IMU 

I/O 

kHz 

MHz 

ms

|XS

ns

NAD

NGVD

rpm

TV

V

W

Explanation

aerial profiling of terrain (system) 

direct current 

direct memory access 

gallium arsenide 

gyroscope 

hertz

inertial measurement unit 

input/output 

kilohertz 

megahertz 

millisecond 

microsecond 

nanosecond

North American Datum 

National Geodetic Vertical Datum 

revolutions per minute 

television 

volt 

watt

INERTIAL INSTRUMENT SYSTEM FOR AERIAL SURVEYING

By RUSSELL H. BROWN, with contributions by WILLIAM H. CHAPMAN, 

WILLIAM F. HANNA, CHARLES E. MONGAN, U.S. Geological Survey, and JOHN W. HURSH,

The Charles Stark Draper Laboratory, Inc.

ABSTRACT

An inertial guidance system for aerial surveying has been developed by The Charles Stark Draper Laboratory, Inc., under contract to 

the U.S. Geological Survey. This prototype system, known as the 

aerial profiling of terrain system, is designed to measure continuously points along an aircraft flight path or a terrain profile to a 

positional accuracy of ±0.5 foot (15 centimeters) vertically and 

±2 feet (61 centimeters) horizontally at the 90-percent reliability 

level. Thus, the system's objective is to accomplish rapidly and accurately from a fixed- or rotary-wing aircraft what traditionally would 

be accomplished from ground-based topographic surveys combined 

with aerial photography and photogrammetry.

The strategy for measuring the terrain profile directly beneath the 

aircraft flight path entails (1) continuous determination of the threecoordinate position of the aircraft, once it departs from its initial 

calibration and alinement point, and (2) continuous measurement of 

the vertical distance between the aircraft in flight and the land 

surface so that the desired terrain profile can be computed directly.

The first, and by far the most complex, part of the strategy is 

achieved by use of an inertial navigator that contains an inertial 

measurement unit, in combination with an eye-safe pulsed laser 

tracker, both mounted on a common base. The inertial measurement 

unit features a gimbaled three-axis inertial platform, stabilized in 

the desired orientation by three high-precision gyroscopes, on which 

is mounted an orthogonal array of three high-precision accelerometers. The gyroscopes provide a three-coordinate scheme of reference 

for measuring the rotational angles through which the aircraft 

moves; the accelerometers provide a continuous measurement of 

specific force acting on the aircraft in three dimensions, from which 

a continuous record of aircraft position can be computed if the Earth's 

gravity field along the flight path is known or modeled. Largely 

because of error buildup in the determination of the vertical coordinate of position, the inertial measurements must be supplemented 

with independent positional or velocity information at approximately 200-second intervals. This requirement is met by a gimbaled 

two-axis laser tracker, which measures distances and directions to a 

select number of ground-based retroreflectors which have been surveyed previously.

The second part of the strategy is achieved using an eye-safe pulsed 

laser profiler, which is, in effect, a high-precision altimeter. Boresighted with this is a color television camera which continuously 

records an image of the flight-path trace over the terrain. An onboard 

computer accepts data from the sensors in the aerial profiling of 

terrain system and performs the computations necessary for calibraUse of brand, firm, and trade names in this report is for identification purposes only and 

does not constitute endorsement by the U.S. Geological Survey.

tion and alinement of the inertial measurement unit, navigation of 

the aircraft, and position and velocity updating of the inertial navigator. The computer logs its results in an onboard magnetic tape 

recorder; selected results are deliverable as well to the sytem operator's keyboard and display unit. Postflight computer processing of the 

recorded data involves such functional steps as editing, renavigating, 

compressing, introducing a refined gravity model, state-space 

Kalman filtering, back filtering, optimal smoothing, and recombining. The results are used to generate a new magnetic tape of 

terrain-profile data in coordinates of latitude, longitude, and elevation, which are time-correlated with a magnetic tape containing the 

color video imagery.

A typical mission is accomplished with the aerial profiling of terrain system installed in a DeHavilland Twin Otter aircraft, flown at 

a speed of 115 miles per hour (105 knots) and at a height above the 

terrain of 2,000 feet (610 meters) over, for example, a 20- by 20-mile 

(32- x 32-kilometer) land area of gently rolling topography. An advance field-survey party sets out and surveys a minimum of three 

retroreflectors in a triangular array; but the number set out may go 

as high as 15, spaced in a way that will allow retroreflector-tracking 

gaps of no more than 300 seconds throughout the intended flight.

Upon completion of a mission and the subsequent postflight data 

processing, the terrain-profile data and video imagery are viewed 

simultaneously on the screen of a video receiver to permit identification and deletion of any unwanted anomalies which might have been 

caused by dense vegetation or manmade structures. Special flight 

missions, restricted to using the aerial profiling of terrain system as 

a means to establish horizontal and vertical control at sites of unsurveyed retroreflectors or as a means to control or monitor the position 

and orientation of other onboard remote sensing instruments, do not 

require operation of the laser profiler.

Performance-evaluation tests of the aerial profiling of terrain system indicate that the ±0.5-foot (15-centimeter) vertical and ±2-foot 

(61-centimeter) horizontal accuracy specifications have been 

achieved. The principal source of error is attributable to uncertainties of gravity anomalies. If future refinements are made for controlling the thermal environment of the aerial profiling of terrain 

system, with special attention to the several electronics packages, 

then gravity-disturbance data may be routinely gleaned from the 

flight-mission data. This, in turn, would permit improvements in the 

currently achieved accuracies. Also, the fully implemented Global 

Positioning System of satellites may offer potential benefits to the 

aerial profiling of terrain system, specifically with regard to providing an additional independent measurement of velocity for the inertial navigator.
INSTRUMENT SYSTEM FOR AERIAL SURVEYING

INTRODUCTION

DEVELOPMENTAL INCENTIVES

The U.S. Geological Survey incentives for the 

instrument-development work described in this paper 

were first (mid-1960's) identified with the hydrologist's 

and cartographer's search for a more efficient, quick, 

and economical way to collect precise terrain-profile 

data at innumerable (tens of thousands) field sites. The 

hydrologist's need related to the ongoing work of defining more precisely the geometry of stream valleys nationwide; the cartographer's need keyed to the perennial quest for improved mapmaking tools and 

techniques.

With the National Flood Insurance Act of 1968, as 

amended by the Flood Disaster Protection Act of 1973, 

the U.S. Congress created the National Flood Insurance Program, which was aimed at improving the 

management of property development in the flood 

plains of streams. The Act required the Secretary of 

Housing and Urban Development to provide each of the 

more than 20,000 flood-prone communities with a detailed flood-insurance study by August 1, 1983. Each 

such study was to include maps showing the aerial 

extent and elevation of the 100-year (1-percent chance 

of occurrence in any given year) flood. The potential 

workload impact of this Congressional mandate upon 

hydrologist and cartographer alike was formidable.

In the early 1970's, hydrologists and geomorphologists were pressing the search for better quantitative 

expressions to relate hillslope profiles and stream gradients to watershed runoff characteristics. Corollary 

interests concerned attempts to expedite geologic mapping in hilly forested terrain, where important geologic 

features were often difficult to spot and trace. The foregoing combination of work and interest, whose progress often was thwarted by a paucity of data on longitudinal and transverse stream-valley profiles, prompted 

the Director in December 1971 to solicit from the several Operating Divisions of the Survey, thoughts on the 

technical feasibility of developing airborne instrumentation to measure terrain profiles. The director followed this invitation with the establishment in March 

1972 of an ad hoc interdivisional committee to explore 

technical feasibility, and that committee's report in 

May 1972 became the stepping stone to the research 

effort described in this report.

DEVELOPMENTAL STRATEGY

Early in 1972, when Brown and Mongan began to 

analyze the potential for developing an airborne instrument system capable of measuring terrain profiles, 

some quantitative and substantial qualitative results

from a variety of earlier flight experiments were already available. Most flights had featured large multiengine aircraft that included an Air Force B-17, a 

civilian Douglas DC-3, and a Navy Lockheed Constellation. In all flights, the laser distance-ranging equipment that was carried was bulky and heavy and 

required a substantial amount of electrical power to 

operate. The limited quantitative results clearly 

snowed that, unless a way were found to define very 

precisely the aircraft path of motion, the overall accuracy in vertical measurements would only approach 

±1 ft (0.3 m) and, in horizontal position, might range 

from tens to hundreds of feet (3-30 m).

Very early in the development analysis, therefore, 

thought processes were conditioned by the following 

rather obvious facts:

1. Aircraft size should be held within the relatively 

light category [less than 12,500 Ib (5,670 kg)] to 

minimize the logistics for long-term continuous 

field use in the Survey.

2. Instrument system should be compact and 

lightweight and should require a minimum of 

electrical power commensurate with the constraints imposed by the choice of aircraft.

3. Instrument system should be capable of defining 

continuously and precisely the aircraft threedimensional path of motion.

4. Instrument system should be self-contained and 

should not require any ground-based electromagnetic radiation (radio or microwave) for its successful operation.

It was fact number 4 that almost immediately 

steered the development analysis into the available instrumentation literature on state-of-the-art developments for stable platforms deployed in satellites and 

inertial guidance systems. The explorations through 

the literature paid off in April 1973 when the first 

schematic was drafted to show what should constitute 

the ideal airborne terrain-profiling system. That schematic featured an inertial navigator to measure continuously the three-dimensional motion of the aircraft, a 

laser distance-measuring device to collect terrainprofile data, and an onboard digital computer to process 

and combine all measured data into the desired formats.

The literature explorations repeatedly associated the 

names of Dr. Charles Stark Draper and the Instrumentation Laboratory of the Massachusetts Institute of 

Technology, which Dr. Draper headed, with some of the 

most significant advances in the technology of inertial 

navigation and guidance and the manned space flights 

to the Moon. A logical and final step, therefore, in the 

developmental strategy for the work described herein 

was to seek Dr. Draper's endorsement of the proposed 

new airborne instrument system. Endorsement in prin-
INTRODUCTION

ciple was obtained in June 1973 in a meeting with 

Dr. Draper and one of his principal staff advisors and 

colleagues, John W. Hursh. Over the next year, the 

foundation was laid for his laboratory, which in the 

interim had become the independent Charles Stark 

Draper Laboratory, Inc., to undertake the necessary 

work through contract with the Survey.

DEVELOPMENTAL HISTORY

In the early 1970's, as national concerns grew over 

property developments in the flood plains of streams, 

the resolve to improve flood-plain management grew 

concurrently. This triggered in the minds of a few Survey hydrologists a desire to formalize a modest beginning research on improved techniques for flood-plain 

delineation. Thus in August 1971, Holland W. Carter 

(ret.), then Chief, Surface Water Branch, and the late 

Roy E. Oltman, then Assistant Hydrologist for Research and Technical Coordination, arranged for the 

assignment of George W. Edelen, Jr., to the research 

problem of investigating and comparing various alternative procedures for flood-plain mapping. This investigation was to feature cost comparisons and was to 

include examination of the prospects for electromagnetic measurements from aircraft.

Edelen's early work coincided with the Director's 

expressed interest in examining the technical feasibility of developing an airborne instrument system for 

surveying. Thus, Edelen served on the ad hoc interdivisional committee appointed to explore the feasibility 

issue. In the wake of that committee's report, Edelen 

arranged with the U.S. Navy Oceanographic Office to 

test fly laser distance-measuring equipment in a Lockheed Constellation aircraft. The intent was to profile 

the terrain along U.S. Highway 60 west of Richmond, 

Va., and to monitor vertical motion of the aircraft with 

an accelerometer; however, severe air turbulence vitiated most results. This served mainly to underscore the 

merit in proceeding with an analytical appraisal of the 

principal features needed in an airborne instrument 

system geared to the Survey's unique field surveying 

and operating requirements. Preliminary work on such 

an appraisal had been started already, but, at this 

point, the effort was intensified with appropriate administrative sanction.

Several early milestones in this developmental history are in the two preceding introductory sections and 

are not repeated here. It is pertinent to note, however, 

that Dr. Draper's June 1973 endorsement in principle 

of the proposed new airborne instrument system, 

coupled with an opportunity to view state-of-the-art 

inertial guidance hardware already tested and proven 

by Draper Laboratory, effectively spurred the earnest 

work to shape the initial research tasks. Thus, at 

widely spaced intervals in the period from June 1973 to

June 1974, several all-day technical discussion sessions were held at Draper Laboratory. Participants included a number of their professionals as well as Survey professionals from the several Operating Divisions. 

The result was a "Request for Proposal" (RFP No. 5507) 

which the Survey delivered to Draper Laboratory on 

April 8, 1974. The RFP described a proposed 6-month 

engineering analysis, or technical feasibility study, 

and, in essence, circumscribed the research task with 

the following specifications:

1. Define the prospects for designing and building an 

airborne instrument system capable of 

a. Determining continuously in real time (virtually 

instantaneously) the three-coordinate position 

of the aircraft.

b. Continuously profiling terrain along the vertical 

trace of flight path.

c. Achieving absolute (referenced to local control) 

position and profile accuracies of ±0.5 ft 

(0.15 m) vertically and ±2 ft (0.61 m) horizontally.

d. Deployment in relatively light fixed- or rotarywing aircraft.

e. Operating at relatively low flight altitudes, arbitrarily defined as below 3,000 ft (914 m) above 

the terrain.

f. Providing vertical pointing stabilization for a 

small camera.

g. Being updated and adjusted in flight with respect 

to its precise orientation and position.

h. Tying aircraft position data to ground-control 

points in local survey area.

2. Define flight and operating techniques for the subject instrument system.

An added stipulation, implicit in the foregoing specifications, limits field use of the proposed airborne instrument system to nonmountainous terrain. Furthermore, the specifications were drawn primarily from 

requirements of the flood-plain delineation problem, 

which entails fitting floods of given magnitudes, such 

as the 100-year flood or the 50-year flood, into the local 

stream-valley geometry. The fitting exercise uses precisely measured valley cross sections at a sequence of 

selected sites and processes the relevant data through 

what a hydrologist terms a "step-backwater computational procedure." In the field-investigation programs 

of the Survey alone, about 10,000 stream-valley areas 

require such flood-plain definition a singular need 

that drives the stakes quite high for success in developing airborne instrumentation that can offer substantial 

work savings in manpower, time, and cost. Reflections 

of this particular Survey need are evident in each of 

several other Federal agency programs.

With the research task defined, the heart of the proposed instrument system was seen to be the inertial
INSTRUMENT SYSTEM FOR AERIAL SURVEYING

navigator which establishes the three-coordinate 

reference-carrying capability. This suggested a versatility in using the system that would transcend the 

rather limited variety of field problems involving terrain profiling. Some expression of the range of field 

uses, identified as time progressed, is given in the section titled "Uses for the Instrument System" and is 

seen to embrace the study programs of all FieldOperating Divisions in the Survey. Furthermore, the 

time and costs projected by Draper Laboratory for the 

necessary contract work clustered principally around 

the inertial navigator. The laser profiler, therefore, 

was recognized, even at this early juncture, as a lowcost and relatively simple adjunct to the heart of the 

instrument system that would enable early flight testing on important problems that involved terrain profiling. Thus, when the need for a system acronym arose, 

project personnel coined "APT" to signify Aerial Profiling of Terrain.

Effective June 24, 1974, the Survey contracted with 

Draper Laboratory (contract no. 14-08-0001-14548) 

for the referenced 6-month engineering analysis, and 

the author was designated Technical Officer for this 

contract. His counterpart at Draper Laboratory, John 

W. Hursh, was designated Principal Investigator. The 

Survey monitored the advancing work regularly and 

closely and stayed involved through technical discussions and field data inputs as questions arose. The 

Draper Laboratory study culminated in a 225-page 

open-file report (Desai and others, 1975).

The significant findings in the report are summarized as follows:

1. An airborne instrument system can be built that 

will perform to the stipulated precision.

2. The airborne system will require update (position 

and velocity vector) information at 3-minute intervals during the flight mission.

3. The update information required by the system can 

be supplied by an onboard tracker component.

4. The inertial and tracker components should be built 

to function as one integral unit.

5. Inflight calibration of the system will require an 

inexpensive retroreflector mounted on each of 

three presurveyed ground-control points. These 

will be tracked, one at a time, by the tracker.

6. The gyroscopes and accelerometers in the system 

should be state-of-the-art quality (Draper Laboratory supplied) if the Survey's high-precision specifications are held; they can be available commercial quality if lower precision specifications can be 

tolerated.

The Survey arranged for outside technical reviews 

by selected scientists in the Harry Diamond Laboratories, part of the Army Materiel Command, in Washington, D.C. Through those Laboratories arrangements

also were made for reviews by scientists at the Army 

Missile Laboratory, Huntsville, Ala. The reviews focused on two fundamental areas: The mathematical 

developments that undergird the concept of a tracker 

device for updating the inertial navigator in flight, and 

the performance capabilities of the Draper Laboratory 

gyroscopes (gyros) and accelerometers in terms of 

meeting the overall precision specified for the complete 

instrument system. The reviews completely supported 

the Draper Laboratory findings.

To anticipate the ultimate need for a decision and 

commitment on when and how to continue contract 

work at Draper Laboratory, a briefing was arranged for 

the Survey Executive Committee, which comprised the 

Director, principal administrators in the Director's office, and the Division Chiefs. The briefing was given on 

December 18,1974, and highlighted work progress and 

outlook. After a followup briefing on May 7, 1975, the 

committee recommended that the contract effort at 

Draper Laboratory be extended through completion of 

the general design for the full instrument system. That 

work phase, estimated to require 18 months, officially 

began June 28, 1975, with approval of the contract 

papers and allocation of an initial sum of money. In 

the hiatus that otherwise would have occurred between 

completion of the engineering analysis and start of the 

general design for the overall instrument system, the 

Survey contracted with Draper Laboratory for a series 

of laboratory experiments with laser devices. Primary 

purpose of the experiments was to accumulate pertinent operational data that would lead to the best design for the laser profiler. In this work, Draper Laboratory was aided by the cooperative consulting efforts of 

personnel in the Electronics Branch of Harry Diamond 

Laboratories. Results of this 8-month work phase are 

documented in the report by Mamon and others (1976).

Concurrent with their recommendation for continuing the Draper Laboratory contract effort, the Executive Committee requested that cost-effectiveness data 

be assembled, based on the potential Survey uses of the 

proposed airborne instrument system. During summer 

1975, those data were assembled and analyzed. Justification was shown for completing the design and venturing into the costly fabrication phase of the 

instrument-system development as budgetary constraints would allow. Highlights of the costeffectiveness analysis were given to the Executive 

Committee in a briefing on November 12, 1975; the 

Survey-envisioned uses for the instrument system 

came from that analysis.

The contract effort progressed through a series of 

distinct work phases, beginning with the described engineering analysis, as shown in figure 1. The figure 

illustrates the sequence and time frames within which
INTRODUCTION

the phases were accomplished and also shows costs of 

the contract work by phase and fiscal year. When completed, each phase was documented in an open-file 

manuscript report which is included in the list of references.

To stay abreast of the evolving contract work and to 

recognize the diversity in potential field applications 

for the proposed airborne instrument system, the Survey chose to monitor progress directly through monthly 

or bimonthly review sessions at Draper Laboratory. 

These sessions commonly ran 1 or 2 days and involved 

the Survey Contract Technical Officer and at least one 

colleague each from the Geologic, National Mapping, 

and Water Resources Divisions, as well as a colleague 

from what was then the Survey's Office of Land Information Analysis. The individuals who have comprised 

that review team throughout most, if not all, of the 

contract work phases shown in figure 1 are Russell H. 

Brown, William H. Chapman, John M. De Noyer, 

William F. Hanna, and Charles E. Mongan. The indicated mix of Survey representation, commonly combined with one or two representatives, most consistently Zoltan G. Sztankay, from the Electronics Branch 

of Harry Diamond Laboratories, brought top-level review expertise to bear in the very relevant scientific 

domains of cartography, geodesy, geology, geophysics, 

hydrology, mathematics, physics, seismology, and sophisticated instrumentation design and development, 

the last involving especially the fields of acoustics, 

electromagnetics, electronics, and optics. Each review 

session featured an informal, but detailed, exchange of 

ideas, identification of problem areas, and discussion of 

the most promising avenues for resolution. On-the-spot

decisions usually were made, thereby forestalling delays that inevitably would have accompanied attempts 

at resolution by correspondence or telephone.

At about the two-thirds point in the 18-month design 

phase of the contract work, Draper Laboratory held a 

design-review session. On June 22 and 23, 1976, 

Draper Laboratory project personnel described their 

progress on the primary facets of the design effort. 

Handout paperbound text and illustration materials 

totaled more than 400 sheets.

In addition to Survey attendance at the review session, representatives from six other Federal agencies 

and from the Geodetic Survey of Canada attended. The 

discussions surrounding the design review reemphasized and strengthened the favorable determinations of 

feasibility that had been developed earlier.

The design phase terminated virtually on schedule 

with submittal of a 401-page open-file report (Draper 

Laboratory, 1977). The report gives the principal general design features for the overall proposed airborne 

instrument system, with significant backup detail to 

support important design choices. Part of this detail for 

the laser profiler evolved from two sets of limited flight 

tests jointly organized by the Survey and Draper Laboratory in September 1976 and April 1977, respectively. 

The flight-test results are in a key report by Youmans 

(1977).

On March 31, 1977, the Survey contracted with 

Draper Laboratory for the design-verification phase of 

the contract effort. The phase was programmed originally for a 10-month period, with the Survey entitled to 

fund the work incrementally. Due to unavoidable budget delays, this phase was not completed until SeptemWORK SCHEDULE

FUNDS 

BY 

FISCAL 

YEARS

WORK TASKS

ENGINEERING 

ANALYSIS

LAB EXPERIMENTS 

LASER DEVICES

DESIGN ENTIRE 

SYSTEM

INTEGRATE, CONFIRM, 

AND CHECK DESIGN

FABRICATE AND 

ASSEMBLE 

ENTIRE SYSTEM

TOTAL COSTS 

(X1000)

1975

±±-

H

19

I I

76

I I

- 

 . .

1

121

in CD 

t

742

T

: I

m*

»

S

1£

I I

FISC

)77

I I

i

fc

;AL>

19

I I

TAP

78

I I

t

o

. .

CD 

h-.

o

1244

-

1533

s

19

I I

79

I I

COt

o

^^^^ M

1829

19

I I

80

I I

0) t

o

10/1/80

^^^^p

3035

19

I I

81

I I

CO 

0

^^^^B

1688

TOTAL 

COSTS 

(X1000)

75

58

1416

2343

6552

10,444

T = TRANSITION QUARTER

FIGURE 1. Fiscal year history of work on U.S. Geological Survey contract by Draper Laboratory, Cambridge, Mass., 1975-81.
INSTRUMENT SYSTEM FOR AERIAL SURVEYING

her 30, 1978. Effective July 13, 1978, Lowell E. Starr 

was named the Survey Contracting Officer's Technical 

Representative, and Russell H. Brown and William H. 

Chapman were named his Technical Designees.

The design-verification phase represented the last 

necessary step before undertaking the fabrication 

phase. Design verification embraced complete integration and checking of all general design features, as well 

as detailed laboratory testing of key components, circuits, and interface electronics. The end products were 

the drawings, assembly and test procedures, and process specifications ready for submittal to the fabrication 

facilities.

The design-verification phase is documented in a 

narrative report published by Draper Laboratory 

(1978). Companion to the narrative report is a 

"Technical Data Package" that contains all approved 

drawings, source control drawings, process specifications, and software documentation. The package is part 

of the official body of open-file information maintained 

at Draper Laboratory, Cambridge, Mass., and at the 

Survey National Center, Reston, Va.

With no break in time, the fabrication phase began 

and spanned the 3 fiscal years from October 1,1978, to 

September 30, 1981 (see fig. 1). In the negotiations for 

this phase, it became evident very quickly that if 

Draper Laboratory fabricated, in the absence of any 

commercial suppliers, the principal inertial components (three gyros and three accelerometers), the contract costs quickly would overrun the Survey's limited 

budget resources. Conferences between the Survey and 

Draper Laboratory elicited an alternate and less costly 

fabrication plan for the prototype APT system, which 

was based on using gyros and accelerometers that were 

emerging as "surplus" to an Air Force contract. In 

adopting this plan, however, some instrument design 

constraints were relaxed, with no relaxation in overall 

precision, to accommodate the existing gyro and accelerometer hardware. The principal relaxed constraint concerned the operating temperature environment for the gyros and accelerometers, and this, 

thereby, invoked a need to build cooling coils into the 

housing of the inertial navigator unit. This, in turn, 

mandated a decision to air condition the aircraft cabin, 

with the end result that the complete prototype APT 

system would need to be fielded in an aircraft larger 

than that envisioned in the Draper Laboratory design 

report. Thus, this first system was designed to be flown, 

through its test, demonstration, and early operational 

phases, in a DeHavilland Twin Otter aircraft rather 

than the earlier envisioned Rockwell Commander 

twin-engine aircraft. The fabrication phase terminated 

on schedule and is documented in a five-volume report 

(Draper Laboratory, 1982).

ACKNOWLEDGMENTS

The authors acknowledge the insights and effective 

support of top managers and scientists in the Survey, 

Harry Diamond Laboratories of the U.S. Army Materiel Command, and Draper Laboratory.

Many individuals and organizations contributed to 

the success of this instrument development project. 

Some are identified at pertinent places in this paper. 

However, in recognition of especially supportive and 

significant contributions, the following individuals, 

and their organizational affiliations, are gratefully 

identified:

U.S. GEOLOGICAL SURVEY 

James R. Balsley, former Assistant Director, Land

Resources (ret.) 1 

Rolland W. Carter, former Regional Hydrologist,

Southeastern Region (ret.) 

Philip Cohen, Chief Hydrologist 

Joseph S. Cragwall, Jr., former Chief Hydrologist

(ret.) 

Edward J. Cyran, Cartographer, National Mapping

Division

John M. De Noyer, former Chief, Earth Resources 

Observation System Program; currently Geologic 

Division

O. Milton Hackett, former Associate Chief Hydrologist (ret.)

Warren W. Hastings, former Assistant Chief Hydrologist, Research and Technical Coordination (ret.) 

Joseph T. Long, former Chief, Branch of Field Surveys, National Mapping Division (ret.) 

Hugh B. Loving, former Deputy Assistant Chief, Research and Technical Standards, National Mapping Division (dec.)

Robert H. Lyddan, former Chief Topographer (ret.) 

Edward A. Moulder, former Assistant Chief Hydrologist, Research and Technical Coordination (dec.) 

Roy E. Oltman, former Assistant Chief Hydrologist,

Research and Technical Coordination (dec.) 

William A. Radlinski, former Associate Director

(ret.)

Richard P. Sheldon, former Chief Geologist (ret.) 

Rupert B. Southard, former Chief, National Mapping

Division (ret.)

Lowell E. Starr, Assistant Chief for Research, National Mapping Division 

HARRY DIAMOND LABORATORIES 

Lyndon S. Cox, Plans and Operations Office 

Robert G. Humphrey, Physicist, Near-Millimeter 

Wave Branch (ret.)

*ret. retired 

dec. deceased 

res. resigned
INSTRUMENT SYSTEM

Joseph M. Kirshner, former Chief, Fluid Systems

Branch (ret.) 

Clarence F. Ravilious, former Mathematician, Fluid

Systems Branch (ret.) 

Helmut Sommer, former Director, Research and

Technology Division (ret.) 

James E. Spates, former Chief, Plans and Operations

Office 

Zoltan G. Sztankay, Chief, Near-Millimeter Wave

Branch

U.S. ARMY MISSILE LABORATORY

Charles A. Halijak, Professor, University of Alabama

Joseph S. Hunter, Aerospace Engineer, Inertial Systems Development

J. V. Johnston, Research Aerospace Engineer, Inertial Systems Development (ret.)

William W. Stripling, Supervising Aerospace Engineer, Inertial Systems Development (ret.)

H. V. White, Aerospace Engineer, Inertial Systems 

Development

THE CHARLES STARK DRAPER LABORATORY

John H. Barker, Project Program Manager

George T. Bernard, Technical Illustrator

William G. Denhard, former Department Head (ret.)

Charles Stark Draper, Senior Scientist

William A. Drohan, former Associate Division

Leader (res.)

Leo F. Hughes, former Section Chief (ret.) 

Irving M. Levin, former Staff Member (ret.) 

Glenn Mamon, Project Technical Director 

Roger E. Medeiros, Staff Member 

Benjamin S. Smith, former Section Chief (ret.) 

J. Arnold Soltz, Section Chief 

Douglas S. Youmans, former Staff Member (res.)

INSTRUMENT SYSTEM

OVERALL SYSTEM CONCEPT

Some important clues have already been given regarding the nature of the overall airborne instrument 

system. From these, the realization should be emerging 

that, if the Survey earth science measurement criteria 

are to be met, the heart of the instrument system, the 

inertial navigator, must be able to determine instantaneously, very precisely, and continuously its position 

and its orientation. The former implies a capability for 

navigation, and the latter, for pointing.

The uniqueness of an inertial navigator is that it 

combines both the navigation and the pointing capabilities in a computer and a mechanism known as an inertial measurement unit (IMU)2 . The IMU is an assemblage of gyros and accelerometers. For it to function as 

it should, it must be able to measure or detect such 

physical vector quantities as gravity (the local vertical) 

and the Earth's rotation. The directions of these vectors are "remembered" by the gyros, which establish 

the needed references for pointing or orienting the 

whole instrument. The manner in which the references 

are created may be likened to an orthogonal set of x, y, 

z coordinate axes fixed on an inner or stable platform 

that actually supports three orthogonally mounted 

gyros. The stable platform is isolated from the disturbing influence of aircraft rotational motion (pitch, 

roll, and yaw) about the three axes by means of a 

gimbal-support structure (see section titled "Gimbal 

System"). The rotational movement of individual gimbals can be offset or compensated for through the actions of their individual torque motors, which, in turn, 

are controlled by signals from the gyros as relayed 

through the appropriate servoamplifiers.

An inertial navigator cannot operate perfectly, 

owing to inescapable built-in mechanical imperfections 

and imprecise knowledge of variations in the Earth's 

gravitational field along its path of motion (the flight 

path). Thus, the position and orientation data yielded 

by the instrument drift slowly away from the truth (the 

correct values). This can be overcome by providing 

high-quality position-data samples periodically from 

an independent source, and the frequency with which 

this is done is related to the desired overall measurement precision. To satisfy the Survey precision criteria, 

the requirement is to update the inertial navigator 

with position data at 3-minute intervals; the independent source for obtaining those data is a laser-ranging 

device called a tracker, mounted in the aircraft in the 

same structural housing as the IMU. During flight, the 

laser tracker locates and locks onto a retroreflector 

target mounted over a known ground-control point. 

While the target is in view (about 30 seconds), a stream 

of range and pointing-angle measurements is collected.

The way in which the inertial navigator and tracker 

complement each other and combine to make an outstandingly high-quality instrument system can best be 

illustrated in the following manner. First, the inertial 

navigator is a device that can very easily recognize and 

measure high-frequency-type accelerations and angular rotations in other words, the normal motions of 

the carrying vehicle (the aircraft). Its performance is 

corrupted, however, by very low frequency angular rotations caused by very slow and random gyro drift and 

other small systematic errors. Second, the tracker is an

2The term "inertial measurement unit" tends to mislead, in the sense that inertia cannot 

be measured directly. The unit simply features mechanical components (gyros and accelerometers) which, in responding to the phenomenon of inertia, deliver the measurements 

needed to define a "specific force" vector (containing gravitational and nongravitational 

accelerations) acting upon the carrying vehicle.
INSTRUMENT SYSTEM FOR AERIAL SURVEYING

instrument that measures best in a way that allows 

good definition of the low-frequency types of rotations 

and disturbances. The tracker has virtually no capability for measuring or defining the high-frequency-type 

disturbances. Finally, for the best performance characteristics to be realized, the inertial navigator and the 

tracker must be optimally combined to obtain a nearideal (minimal total error) measurement capability 

over the full frequency range for the expected accelerations, angular rotations, and disturbances. These several points are diagrammed schematically in figure 2.

An important attribute of the inertial navigator is 

that, in fulfilling its navigation capability, its IMU 

component directly senses or measures accelerations of 

the carrying vehicle, and these are then mathematically integrated once to yield velocities and twice to 

yield travel distances. Because integration is a smoothing process, any inherent measurement errors are 

greatly reduced; thus, the inertial navigator is especially good at precisely resolving high-frequency types 

of motion.

Implicit in the foregoing discussion is a capability for 

processing the sensed or measured data and performing computations; thus, an airborne digital computer is 

indeed a central component in the complete instrument 

system. Although hardware and related software details are described in the section titled "Computer," a 

few advance observations are pertinent here.

The airborne computer and its accessories must be 

versatile enough to perform a variety of tasks that 

include accepting or logging the sensed and measured 

data from the IMU, tracker, and any ancillary earth 

science measuring equipment; processing the accepted 

data and performing necessary computations; storing 

and displaying the appropriate data; and accomplishing all of this in accordance with suitably developed 

instructions (software). Selection of the most appropriate computer must reflect consideration not only of the 

foregoing tasks but also the need for some reserve ca-

_

< ill O

KOC OC OD oc i- <n uj

aFAVORS HIGH-FREQUENCY MEASUREMENT DATA 

"FAVORS LOW-FREQUENCY MEASUREMENT DATA

FIGURE 2. Frequency response characteristics of inertial navigator and tracker to vehicle motion.

pacity in processing and computing speeds, memory 

(storage) size, and input and output capabilities. This 

represents a deliberate choice to build into the instrument system greater flexibility for planning and executing a larger variety of data-gathering flight missions.

Any sincere attempt to detail the instrument system 

leads inescapably to and through descriptions of the 

principal components. Because their introduction to 

data-gathering programs in the earth sciences is relatively new, they deserve more than casual attention. In 

this paper, therefore, each principal component in the 

overall instrument system is described in detail with 

pertinent extracts from the supporting body of scientific principles and theory.

INERTIAL NAVIGATOR

The earth sciences community is gradually accumulating exposure to and experience with inertial-type 

measuring instruments with the advent of commercially available portable and land-vehicle-mounted 

units used in determining geographic position. The 

precision requirements for such units, however, have 

been at least an order of magnitude coarser than those 

specified by the Survey for the instrument system 

herein described. Furthermore, the Survey specifications require deployment in a light aircraft, rather 

than in land vehicles.

To comprehend the significance of an inertial navigator, one needs to reexamine, briefly, Newton's three 

basic laws of motion and the commonly used term, 

"inertia." These laws of motion feature three principles 

which have proved valid for all mechanical problems 

not involving speeds comparable to the speed of light 

and not involving atomic or subatomic particles. The 

first law, also know as Galileo's law of inertia, states 

that a particle not subjected to external forces remains 

at rest or moves with constant speed in a straight line; 

the nongravitational measurements of accelerometers 

contained within an inertial navigator are referenced 

to an inertial coordinate system in which a body moves 

with constant velocity as long as no force acts on it. The 

second law states that the acceleration of a particle is 

directly proportional to the resultant external force 

acting on the particle and is inversely proportional to 

the mass of the particle; a more general statement of 

this law is that the force equals the time rate of change 

of linear momentum, where momentum is the product 

of the mass and the velocity of the particle. The more 

general statement reduces to the more specific one if 

mass does not change with time, as is the case with an 

inertial navigator. The more general statement involving momentum also provides a reference for defining 

inertia that property of matter which manifests itself 

as a resistance to any change in the momentum of a
INSTRUMENT SYSTEM

body. The third law states that, if two particles interact, the force exerted by the first particle on the second 

is equal in magnitude but opposite in direction to the 

force exerted by the second particle on the first. This 

law explains why the contact force exerted by the support structure on the triad of accelerometers is the inertial reaction to the vehicle acceleration and is, therefore, opposite in direction to vehicle motion.

The literature on instrumentation for inertial guidance and navigation contains many definitions of inertial navigators. Inasmuch as no single definition seems 

adequate for the present purpose, the following composite was collected piecemeal with interpretations 

from several inertial-navigation textbooks (see 

"Selected References"):

"An inertial navigator is a self-contained instrument package designed to track continuously the 

velocity and position of the vehicle that carries it. 

Its proper functioning requires no external electromagnetic radiation. It consists of accelerometers 

that measure a combination of inertially referenced specific force (force per unit mass) and gravitational force; gyros that provide the inertially 

referenced three-coordinate scheme by virtue of 

precessing (spin axis slowly changing direction) in 

reaction to disturbing torques or vehicle rotations 

at right angles to the precessions; an assumed 

model of gravitational force to permit extraction of 

inertially referenced specific force from accelerometer measurements; a computer to integrate with 

respect to time the specific force data, to obtain the 

navigator velocity and position; an updating device to correct or compensate for errors in navigator velocity and position caused by time-dependent 

drifts of gyros and other instrument components; 

and housing materials to interconnect components 

and to stabilize and regulate the physical environment of the instrument package." 

The specific force, equivalent to force per unit mass, 

may be seen to have units of acceleration by expressing 

Newton's second law of motion in the simple form of 

force (F) equals mass (M) times acceleration (a) 

F = Ma and by subsequently dividing both sides of 

the equation by the mass. As previously noted, the 

measured specific force is opposite in direction to vehicle motion as a consequence of Newton's third law of 

motion.

The variety of Survey field applications envisioned 

for the APT system contains the common element of 

field surveys of some kind in a relatively localized geographic area. This prompted selection of a geographic 

coordinate system of latitude and longitude, based on a 

reference ellipsoid and specifically upon the North 

American Datum of 1927 (NAD of 1927), and elevation, 

based on a geopotential surface termed the "geoid"

and specifically on the National Geodetic Vertical Datum of 1929 (NGVD of 1929). This Earth-fixed coordinate system serves as the reference for all fully processed position data. Other coordinate systems are 

used, however, in implementing the inertial navigator; 

for example, the stable platform established by the 

three gyros in the APT system is mechanized in a 

tangent-plane orientation, convenient for inertial navigation when individual flight missions will operate 

over relatively small areas of the Earth areas, for 

example, that have a radius of less than 200 mi (322 

km). This means that the x-y plane of the platform is 

maintained normal (and the z axis parallel) to the direction of the actual plumb line at the point where the 

system is calibrated and alined before a flight mission. 

To maintain this initial tangent-plane orientation 

throughout the flight mission, the system need only 

compensate for Earth rotation.

The tangent-plane coordinate system was judged to 

be superior to a locally level coordinate system because, 

in the latter system, knowledge of changing torque 

rates required to maintain the platform in the system 

could not be recovered for postsurvey data processing 

(postprocessing). The locally level coordinate system, 

however, is used as a reference for the processed velocity data, having been transformed to this system from 

the tangent-plane system. In the locally level system, 

the x-y plane of the stable platform is maintained tangent to the surface of a mathematically defined reference ellipsoid, and the z axis is maintained parallel to 

the ellipsoid normal. Gyrocompassing permits the xaxis to be oriented along geographic north and the yaxis along geographic east. The direction of the z-axis 

departs from the direction of actual gravity approximately by a quantity known as the deflection of the 

vertical.

The requirement for a model of the gravitational 

field is especially important where an inertial navigator is operating in a near-Earth environment, as in an 

aircraft, and where precise knowledge of the elevation 

position coordinate is required. Both conditions apply 

to the present system. Much experimentation will be 

warranted in determining how accurate a gravity 

model (gravitation plus centripetal acceleration associated with Earth rotation) is required for a given spacing of ground-surveyed control points to achieve a 

stated precision in vertical and horizontal position coordinates.

Inertial Measurement Unit

The principal physical component of a threecoordinate inertial navigator, such as that designed for 

the APT system, is the IMU composed of three gyros 

together with appropriate gimbals (supporting frames)
10 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

and servo-drive motors to establish a stable platform 

that will hold the specified three-coordinate scheme of 

reference in the correct orientation, and three accelerometers mounted on that stable platform to measure the resultant of the specific forces referenced to 

each coordinate axis. Beyond these major items, the 

IMU includes a number of subsidiary items, such as 

electronics packages, pickoff and readout devices to accomplish the required measurements, and all the appropriate wiring and mechanical linkages.

An IMU can be visualized as a key component in the 

inertial navigator, which is designed to solve problems 

in geometry, inasmuch as the ultimate task is to locate 

points along its path of motion and to reference them 

very precisely (in space and time) to some specified 

coordinate system. The challenge, then, is to design 

and build a collection of physical components, the heart 

of which is labeled an IMU, capable of simulating the 

specified coordinate system, measuring the resultant of 

all applied forces (including gravity) referenced to that 

coordinate system, and processing that information in 

real time (or after the flight mission) into the desired 

position and orientation or attitude information.

An important key in guaranteeing inertial navigator 

performance that is dependable, predictable, and at the 

designed level of precision is the provision of a benign 

thermal environment for the inertial components; that 

is, the gyros and accelerometers. The goal is to simulate the stable and perfectly controlled conditions of the 

laboratory throughout the expected vagaries of the 

field-operating environment. In a number of missile or 

astronautical applications, this has required adjunct 

large and heavy refrigeration units which might have 

compromised the Survey stipulation that the APT system be capable of deployment in a relatively light aircraft. Thus, in the original design, Draper Laboratory 

designers chose a high working-level temperature inside the IMU of 128°F (53°C) stabilized to within 

±0.1°F. In such a design, air would be the heat transfer 

medium, and aircraft cabin air, the ultimate heat sink. 

Thus, under most expected operating conditions, heat 

would need to be supplied the IMU via built-in electrical heaters. Fabrication costs for this high-temperature 

IMU precluded its use in the prototype system, but the 

concept is sound and should be implemented for any 

future operational system. Departures from the 

original-design temperatures are described at the end 

of the section on temperature control.

Gyroscopes

Fundamental to the successful implementation of 

the IMU are the gyros that provide the physical means 

for creating a stable platform to carry the given coordinate scheme of reference in a specified orientation in 

three-dimensional space. Popular qualitative concepts,

harking back all the way to childhood toys, convey 

some of the pertinent characteristics of these intriguing devices. However, a more exact and quantitative 

detailing of selected gyro properties is germane to the 

present description.

The pertinent nomenclature relating to gyros, displayed with the elemental schematic in figure 3, is 

important to comprehend and properly define significant aspects of their performance characteristics. The 

heart of a gyro is a rotor or wheel free to spin on an axis. 

The spin axis is held in nearly frictionless bearings 

retained in a gimbal which itself is held in low-friction 

bearings retained by the overall supporting structure 

or base; for example, an aircraft. Among the design 

goals for a spinning gyro is a high value for its angular 

momentum along with small physical size. A specified 

angular momentum, therefore, compatible with small 

size, is obtained by concentrating the weight at the 

perimeter and spinning the wheel at a high speed; for 

example, 24,000 revolutions per minute (rpm) is used 

in the APT system and is commonplace in commercially available gyros. The APT system gyros also will 

include a viscous damper placed on one gimbal axis, 

whose function will be described later in this section.

The spin axis of a gyro will maintain its initially 

given orientation in space. The degree of steadfastness 

with which the spin axis will hold its orientation is 

related directly to the magnitude and constancy of the 

angular momentum of the spinning wheel. Furthermore, if the spin axis experiences a rotational movement, causing the direction of that axis to change, it 

will react in a new rotational movement (directional 

change) orthogonal to the first. This gyro property can 

be illustrated by imagining that the inner gimbal supporting the wheel (fig. 3) undergoes a slight rotational 

movement clockwise around the middle axis. The foregoing movement can be caused by applying a parallel

GYRO ROTOR OR WHEEL

OUTER AXIS ("^ INNERGIMBAL

MIDDLE AXIS

OUTER GIMBAL

PRECESSION

FIGURE 3. Gyro and gimbal-support structure.
INSTRUMENT SYSTEM 11

pair of equal and opposite forces to the inner gimbal as 

shown in figure 3. The consequent spin-axis movement 

can be seen to result in its changing its line of direction 

in the plane of the two forces and the reaction or response is a predictable and new spin-axis rotational 

movement, or precession, in an orthogonal plane (fig. 

3). The foregoing pair of applied forces is termed a 

"couple," and the quantitative measure of its capability 

for producing rotation is termed a "torque," which is 

simply the product of one force times the perpendicular 

distance between the pair (the arm in fig. 3).

The above statement is general and not limited to the 

particular plane that passes through the applied forces 

shown in figure 3. The equation that quantifies the 

above relation is called the Law of Gyroscopics, which 

may be expressed in the convenient form,

[Torque] = [(Wheel Inertia) (Spin Velocity)] [Precession Rate]3 .

(N-m) (m2kg) (rad/s) (rad/s)

Shown parenthetically below each quantity are the SI 

(International System) units in which it is commonly 

expressed (see also list of "Symbols and Dimensions"). 

Because the product of wheel inertia and spin velocity 

is angular momentum, a simpler, more convenient 

form of the equation is

[Torque] = [Angular Momentum] [Precession Rate]. (1)

(m2kg/s2) (m2kg/s) (rad/s)

Both of the above equations relate the indicated terms 

in magnitudes only.

Equation 1 is extremely useful in a two-way sense in 

that, for a given or applied torque (magnitude and direction), a particular rate and direction of precession is 

determinable, and, conversely, if a specific rate and 

direction of precession is observed or created, then the 

magnitude and direction of the particular related 

torque is determinable. The principle actually is used 

in both senses as the APT system operates.

The described performance characteristics immediately suggest the possibility for using gyros to set up 

the desired three-coordinate-reference axis scheme. 

However, if the spin axis is to be given a proper chance 

to fulfill its function of remaining stable in a given 

orientation, then it must be isolated in carefully predetermined ways from the several possible rotational motions of the overall supporting structure or base.

In reference to figure 4, note the given orthogonal set 

of three axes comprising the spin axis, the output (gim3The rotational equivalent of F = Ma (which describes translational motion) is T_ = la, 

where T, I, and a are the torque (a pseudovector), the moment of inertia, and the angular 

acceleration (a pseudovector), respectively.

STABILIZED REFERENCE LINE

GYROGIMBALGYRO ROTOR OR WHEEL

INPUT OR SENSITIVE AXIS' 

(QUADRATURE AXIS)

FIGURE 4. Single-degree-of-freedom gyro.

bal) axis, and the input (sensitive) axis, the last of 

which, by definition, is at right angles (in quadrature) 

to the other two. Compared to figure 3, the outer gimbal has been removed, and a damper added to the output axis. Next consider that if the base is moved along 

any arbitrary path so that its original orientation in 

three-dimensional space is changed, then the progressive nature of that orientation change can be fully described in terms of three component rotational rates 

about the foregoing set of orthogonal axes. It is appropriate now to examine individually these three possible 

component rates.

1. Any base-rotation component around the spin axis 

obviously has no effect on changing the direction 

of that axis.

2. Any base-rotation component around the input axis 

(fig. 4) inevitably causes the spin axis direction to 

experience the same rotational rate, and the reaction, as may be predicted from the preceding discussion, is a precession torque which induces rotation of the gimbal and wheel assemblage around 

the gimbal (output) axis. This torque is countered 

by an opposing torque from the viscous damper 

proportional to the angular rate of the gimbal 

around the output axis, which, in effect, integrates the precession torque and affords a way of 

precisely indicating the angular rotation of the 

base around the input axis. An electromagnetic 

device known as a signal generator is used to 

measure the angle between gimbal and base as an 

output voltage.

3. Any base-rotation component around the output 

(gimbal) axis is countered by a resisting torque in 

the viscous damper. This has the effect of tending 

to induce rotation of the gimbal and wheel assemblage around the gimbal axis, leading to the 

predictable precession torque around the input 

axis. Due to the geometry of the gimbal structure
12 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

(see fig. 4), however, the only result is a tendency 

to "cock" the gimbal in its bearings.

Because the gyro spin axis is free to move in response 

to a torque around only one axis, the input axis, it is 

termed a "single-degree-of-freedom" gyro and is so labeled in figure 4. Thus, the gyro measures base rotation 

around the input axis (sometimes called the sensitive 

axis) only and quantifies this as an output voltage proportional to input angle.

The conclusion can now be made that if, for example, 

the input axis in figure 4 is chosen to simulate the local 

vertical coordinate axis and if two similar singledegree-of-freedom gyros are mounted on the same base 

with their input axes oriented orthogonal to each other 

in the local horizontal plane, to simulate stabilized reference lines in the east and true north directions, then 

a stable platform reference scheme in Cartesian coordinates has been instrumented, capable of sensing base 

rotation around any axis. The balance of the instrumentation challenge stems from the need to provide the 

platform with motion-isolating gimbals driven by 

torque motors powered by servoamplifiers receiving 

signals from the gyros. Each gyro signals the angular 

amount it has been forced to deviate because of the base 

motion. This information is resolved along the gimbal 

axes and then is fed to the respective servo-drive or 

torque motors to rotate the gimbals so as to restore 

each gyro to its required stabilized position. In effect, 

therefore, the gyros mutually act to isolate themselves 

from the three component base rotations caused by any 

motion of the carrying vechicle. The entire foregoing 

cycle of events occurs virtually instantaneously, and it 

is repeated continuously.

The particular gyros selected for the Survey instrument system are state-of-the-art hardware fabricated 

by Draper Laboratory. In the eventual production 

model, the gyros are to be fabricated in accord with a 

design that uniquely fits the Survey specifications and 

field-operating environment. In the APT system the 

gyro will carry the Draper Laboratory designation "APTG."

The gyro unit, weighing 1.0 Ib (0.45 kg), is housed in 

a cylindrical case approximately 3 in. (76 mm) long by 

2 in. (51 mm) in diameter. Its principal features are 

labeled in the cutaway view of figure 5. The gyro rotor, 

or wheel, is symmetrical and is supported by opposed 

hemispherical gas bearings. In operation, the rotor 

spins at a constant speed (24,000 rpm) which is sustained by torque from a pair of hysteresis-ring assemblies on the rotor. These ring assemblies are driven by 

a pair of AC motor stators. The complete motor, wheel, 

and bearing assembly is enclosed in a hermetically 

sealed cylindrical "float" (shown in red in fig. 5) which 

is filled with hydrogen. The hydrogen provides gasbearing support and also serves as a thermalconducting medium helping to produce an even temperature distribution over all inside parts of the float 

shell. The float is suspended in a support structure 

(shown in blue in fig. 5) with no mechanical contact, 

except for flexible power leads. The support techniques 

involve electromagnetic centering and a highly viscous 

liquid for flotation. Temperature of the liquid is maintained at a constant 139°F (59°C), to a tolerance of 

0.001°F (0.0006°C), partly through the device of a thermal control heater (see fig. 5) and partly through stainless steel thermal shims used with the center-flange 

mount.

The power leads, which carry electrical power to the 

wheel-drive motor, are flexible arched metallic leads 

that are submerged in the flotation fluid. Semiflotation 

of the leads in the viscous fluid reduces the distorting 

effects that occur when the gyro is subjected to high 

linear accelerations. The leads have their outer ends 

mounted on a baffle plate that has small clearance 

grooves for each lead. These grooves are so small that 

solidification of the fluid at low temperature cannot 

produce shearing forces large enough to damage the 

leads.

In further reference to figure 5, an electromagnetic 

device with inner and outer concentric magnetic circuits is located at each end of the instrument. The concentric rotor assemblies are mounted on the ends of the 

float, and their respective concentric stators are 

mounted in the cylindrical housing. The outer circuit in 

one end features a signal generator that detects the 

rotation of the float with respect to the housing and 

produces a polarized voltage signal proportional to the 

angle of rotation. The outer circuit in the other end 

features a torque generator that is the means for applying precession torques to the float. The inner circuits on 

both ends feature magnetic suspension units to provide 

axial and radial centering of the float pivot within the 

clearance spaces and thereby eliminate mechanical 

pivot-to-jewel friction.

The fluid in the gyro unit supports the float so that 

the residual weight carried by the magnetic suspension, which acts to center the float, is no more than a 

small fraction of a gram under a pull equal to one times 

the Earth's gravitational force. Under conditions of 

shock, vibration, and linear acceleration, bouyant reactions proportional to the applied specific force are generated by the fluid. This means that the acceleration 

reaction force tending to move the float within the gyro 

case depends only on the residual unfloated mass and 

not on the total mass of the float. Furthermore, the 

float velocity generated by the residual unfloated mass 

being acted on by an inertial reaction force is so limited 

by the viscous reaction forces of the damping (flotation) 

fluid that a time period of 30 minutes to 1 hour is 

required for the gimbal to move radially 0.0001 in. 

(0.0025 mm) under a force equal to the force of gravity
INSTRUMENT SYSTEM 13
14 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

near the Earth's surface. Thus, the combination of flotation, magnetic centering, and high damping makes it 

possible to keep the pivots from making contact with 

their bearings even under the most severe environmental conditions of shock, vibration, and acceleration.

Acceleration-compensated bellows are located at 

each end of the gyro. These permit thermal expansion 

of the fluid and are designed to prevent fluid-float mass 

motion under acceleration along the output axis.

Viscous shear forces acting across the damping gap 

between the case and the float provide a near-perfect 

velocity-proportional restraint torque to oppose the 

precession torque, which produces the integrating action that characterizes the gyro unit. The high viscous 

shear torque causes the gyro to behave as a linear, or 

first-order, system with a characteristic time of about 

0.74 millisecond (ms). In the mathematical comparison 

of linear systems, the time characteristic arbitrarily 

represents the time interval over which the transient 

phase of the system response to a stimulus reaches 

63-percent completion. The low value for this gyro signifies a built-in capability for rapidly and precisely 

tracking all stimuli of short duration in a band 

bounded by an upper frequency limit of about 1,400 

hertz (Hz).

For a given angular momentum, the viscosity of the 

fluid controls the ratio of angle input to float-angle 

output, or sensitivity, of the gyro unit. Although a preselected sensitivity value is not usually among the design criteria for a gyro, the value must be determined 

once the gyro has been built to document pertinent 

performance characteristics. For the Draper Laboratory gyro and a wheel speed of 24,000 rpm, the sensitivity is about 0.9.

No matter how carefully a gyro is designed and constructed, it can never be an absolutely perfect device. 

Thus, the spin axis (stabilized reference line) is subject 

to stray torques brought about, for example, by minute 

friction in the support bearings or in the angular displacement pickoff devices. These torques, which develop randomly, cause the spin axis to precess or wander (change direction) in random amounts; this is 

termed "gyro drift." For a given gyro, this random drift 

can be observed and defined through laboratory tests 

and calibration procedures so that, within carefully 

specified limits, proper compensation can be introduced 

when the gyro is put into operating service.

A measure of the Draper Laboratory gyroperformance capability with respect to drift can be 

drawn from a practical comparison with the gyros now 

extant in inertial navigation units used by commercial 

airliners on the long transoceanic international flights. 

These units commonly accumulate gyro-drift errors in 

position of about 1 nautical mile (nmi) per flight hour. 

On a great circle route over the Earth's surface, 1 nmi

[about 6,080 ft (exactly 1,852 m)] is equivalent to 

1 minute of arc. The gyro drift, then, can be stated as 

1 minute of arc per hour. The Earth, however, turns at 

a rate of 15°, or 900 minutes of arc, per hour. Thus, the 

observed drift rate for gyros now in commercial use is 

seen to be about one-thousandth (1 x 10~3 ) of the 

Earth's turning rate. Draper Laboratory personnel 

have adopted this milli-earth-rate unit (meru) as a convenient measurement unit to describe and compare the 

performance characteristics of inertial navigators.

In contrast to commercial gyro performance, the design objectives for the Draper Laboratory gyro require 

holding the accumulation of position errors related to 

gyro drift to about 100 ft (30.5 m) per day or 4 ft (1.2 m) 

per hour. If this distance is recognized as something 

less than 1 x 10~3 nmi,then the gyro drift rate is obviously less than 1 x 10~3 minute of arc per hour, or less 

than one-millionth (1 x 10~6) of the Earth's turning 

rate. In the APT system, therefore, gyro performance is 

projected to be several orders of magnitude better than 

present commercial gyros.

To give some quantitative feel for the relation between gyro drift rate and a causative torque, numbers 

can be substituted in the basic gyro equation (see eq 1) 

to reflect specific performance properties of the Draper 

Laboratory gyro. Equation 1 can be written more formally as

(2)

where JT0 = torque around the gyro output axis in newton meters per radian (N-m/rad) 

L = angular momentum of the gyro wheel in

newton meter seconds (N-ms) 

wj = angular velocity around the gyro input 

axis in radians per second (rad/s).

If the gyro spin axis changes direction (drifts) at an 

angular rate of 1 meru, this means wj = 7.27 x 10~8 

rad/s. For the Draper Laboratory gyro unit, L is approximately 3.75 x 10~2 m2kg/s. Substituting these values 

in equation 2 results in a T0 value of 2.7 x 10~9 m2kg/s2 

N-m for the torque created at the gyro output axis.

To reduce the effects of anisoelasticity. the gimbal (or 

float structure) is made of beryllium, which combines 

an extremely high modulus of elasticity with a low 

density. Also, the properties of the gas bearing have 

been optimized for isoelastic stiffness. To provide isolation from magnetic fields, the entire gyro is surrounded 

by a two-layered, 90°-overlap, metallic shield of high 

magnetic permeability.

The coordinate system of the stable platform is conveniently established by mounting three gyros so that 

their input axes are mutually orthogonal, as shown
INSTRUMENT SYSTEM 15

schematically in the upper part of figure 6. The orthogonal orientation is justified for two reasons: The input 

axes correspond to the x, y, and z axes of a Cartesian 

coordinate system, and this orientation permits minimizing unwanted torques on the gyro float associated 

with any built-in unbalance. The unwanted torques, 

any one of which is at a minimum when forces acting on 

the gyro are oriented along its output axis, are minimized for the three-gyro set by requiring two of the 

output axes to be alined along or near the actual gravity vector acting on the APT system.

Because their thermal sensitivity is greater than 

that of other associated components, the three gyros 

and their required electronics are arranged in their 

own physically separate section of the stable platform. 

This makes it easier to provide a benign thermal environment for the gyros which ensures achieving the 

highest possible performance levels. Uniform temperature distribution in the most thermally sensitive part 

of the gyro the float containing the spinning wheel 

is enhanced by adopting a center-flange mount (see 

fig. 5) for the cylindrical case. The mechanical and 

thermal design features provide added insurance for 

uniform temperature distribution by specifying a good 

thermal conductor from the center-flange mount to the 

gyro ends, and a heater and sensor for each end of the 

conductor so that thermal disturbances at the mount do 

not affect the gyro.

AZIMUTH AXIS

(NORTH)

THREE-GYRO ARRAY

ACCELEROMETER 

TRIAD

EXPLANATION 

SA =SPIN AXIS 

IA = INPUT AXIS 

OA = OUTPUT AXIS

FIGURE 6. Gyro and accelerometer arrays for aerial 

of terrain system.

profiling

Accelerometers

In the hierarchal construction of an IMU, gyros have 

been described as the basic or fundamental devices 

needed to create a stable platform. Not only are the 

changes in attitude of the carrying vehicle, as it progresses along some path of motion, measured and referenced to this platform, but also other instruments can 

be placed on it to measure the resultant offerees acting 

upon that vehicle. Accelerometers are the particular 

instruments used in the next hierarchal construction 

stage for an IMU, and these precision devices accomplish the needed specific-force measurements. A quantitative detailing of selected accelerometer properties, 

along with related excerpts from the classical mechanics, is appropriate in this limited discussion.

The simplest variant of an accelerometer can be pictured (fig. 7A) as a finite mass, M, supported on a 

horizontal frictionless slide bearing or base and 

restrained horizontally by a spring that obeys Hooke's 

law.4 If this accelerometer device experiences an acceleration, a, in a horizontal direction as shown, then the 

spring must supply a tensile force, F, to make the mass, 

M, move along with the base. Through precalibration, 

the spring deflection is a measure of that force. Thus, 

this simple accelerometer measures F in the general 

expression F = Ma which leads to a solution for "a". 

The foregoing solution presumes that no component of 

the acceleration is attributable to gravity. If held stationary, the device can be used to measure the acceleration of gravity by alining its base with the local vertical.

Among the more elegant and sophisticated variants 

now in accelerometer devices, the designs particularly 

germane to this discussion are those that capitalize, in 

part, on the properties of pendulums. The necessary 

fundamentals can be drawn from the schematic of figure IB in which a simple pendulum of the indicated 

length and mass is suspended from a pivot point, P. If 

the pivot point experiences a constant acceleration, a, 

to the right, as shown, then inertia will cause the mass 

to lag behind until the pendulum arm makes the determinable angle, 6, with the vertical. Viewed from P, it 

appears that the pendulum has been subjected to a 

torque, T, whose magnitude is given by the simple 

expression

where k is a constant relating to the physical characteristics of the system. The key to a rigorous mathematical analysis leading to the foregoing expression is 

given by Landau and Lifshitz (1976, p. 11).

4The extension, x, of a spring that is subjected to a tensile force, F, is directly and linearly 

proportional to that force; that is, F = -kx, where k is a constant property of the spring.
16 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

If an equal, but opposite, torque is applied to the 

pendulum in the foregoing situation, then the pendulum will be returned to its undeflected, or "null," posiPRECALIBRATED 

SCALE OF FORCE, F

ACCELERATION, a

ELASTIC SPRING 

RESTRAINT

ELEMENT OF MASS, M 

FRICTIONLESS-SLIDE BEARING

NEWTON'S SECOND LAW: F = Ma

SIMPLE 

ACCELEROMETER

ARM LENGTH, L

ELEMENT OF MASS, M

PIVOT POINT, P

ACCELERATION, a

8

SIMPLE PENDULUM

FOR SENSING ACCELERATION

TORQUE MOTOR

PIVOT POINT, P

ACCELERATION, a 

*

FEED-BACK 

LOOP

ELEMENT OF MASS, M

NULLED PENDULUM

FOR SENSING ACCELERATION

FIGURE 7. Acceleration-measuring devices.

tion. This encourages, therefore, a slight modification 

of the simple pendulum in the manner shown schematically in figure 1C. The intent now is to monitor with, 

for example, a beam of light and the proper photocells 

any incipient movement of the pendulum as acceleration occurs. Then, through a correctly designed feedback loop, an electromagnetically generated countervailing torque can be applied in just the magnitude 

needed to keep the pendulum in its null position. Thus, 

a precalibrated torque magnitude becomes the unique 

measure of an acceleration, and the desired data can be 

read out electrically. Within the limits of the magnetization characteristics of the torque motor material 

and current supply, the torque magnitude tracks the 

changing values of the acceleration. The stability and 

linearity of such torquing devices, however, leave 

much to be desired.

One significant refinement of the foregoing evolutionary review of selected acceleration-measuring 

devices carries the analysis to the particular type of 

accelerometer featured in the Survey instrument system. That refinement capitalizes on the gyro property 

that so precisely relates torque around one axis to angular precession rate around an orthogonal axis. In the 

discussion on gyros, the point that this property provides the means for very precisely generating a small 

torque was defined clearly. Such a torque can be used 

to restore the foregoing pendulum (fig. 7C) to null.

A cutaway and schematic view of the accelerometer 

selected for the APT instrument system is shown in 

figure 8. Note that it features a floated gyro similar to 

that shown in figure 5 but with one important difference a pendulous mass built into the positive end of 

the spin axis. This type of accelerometer may thus be 

regarded as a floated pendulum with a gyro inside, 

magnetically suspended in the radial and axial directions with an electromagnetic device which can generate an electrical signal. As the pendulous mass (fig. 8) 

attempts to move downward under the influence of 

gravity, it tilts the float and, thereby, invokes an output from the signal generator proportional to the angle 

through which the float turns.

Note also that the red-shaded structure, including 

the float component, is free to turn around a vertical 

axis which is also the input axis of the gyro. If rotation 

should occur around that axis, then it will produce a 

consequent gyro precession torque that will counteract 

the tilt of the float. An external servoamplifier and a 

servomotor built into the accelerometer housing or case 

are the means for providing this rotation. By feeding 

the output of the signal generator, with the proper 

phase relation, into the servoamplifier, a closedloop system is brought into operation wherein the precession torque of the gyro restores the pendulum float 

to the null position and keeps it there under varying
INSTRUMENT SYSTEM 17

conditions of acceleration. The formal name given to 

this measuring instrument is "pendulous integrating 

gyro accelerometer."

From the preceding discussion, note that, by observing the rotation rate of the red-shaded structure in 

figure 8 (servo-driven member), the consequent precession torque that is created around the gyro output axis 

is precisely determinable and is related directly to the 

specific force that was imposed along the gyro input 

axis. If, however, the total accumulated angle of rotation is observed instead of the rotation rate, then this is 

equivalent to performing the first time-integration of 

specific force, and, therefore, a precise measurement of 

integrated specific force is obtained, which has the

units of velocity. (Recall that the time rate of change of 

velocity is equivalent to acceleration.) The resolver 

shown in the figure measures that angle, and the measurement data, which comprise the accelerometer output, are fed to the airborne computer. By using three of 

these accelerometers, oriented so that their input axes 

are mutually orthogonal, it becomes possible to measure the specific force on the vehicle carrying the IMU 

along any conceivable three-dimensional path of motion.

For the Survey APT instrument system, the accelerometer will carry the Draper Laboratory designation "APTA." The actual accelerometer unit is as 

shown in the artist's cutaway view of figure 9, and the

AC< JURAT 10*4 4

SfRVO AMPLIMtR AND 

RfSOLVfH READOUT 

f Lf CTHONICS

SERVO MOTOR 

ROTATION

fLOATANGLE

liVROSCOPIC TORQUE

**k SPIN AXIS NULL POSITION 

T ^^ SPIN AXIS

OYRO OUTPUT AXIS

PENOUl COS MASS

FIGURE 8. The single-axis pendulous integrating gyro accelerometer.
18 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

principal features are as labeled. The pendulous integrating gyro component is shown in red, and the servodriven member is in blue.

The unit features a gyro wheel driven at 24,000 rpm 

and supported by hydrogen-filled gas bearings. This 

provides an angular momentum of 2 x 10~4 N-ms 

under a constant [(±0.1°F)(±0.06°C)] operating temperature of 139°F (59°C). The neutrally bouyant and 

symmetrical float which houses the gyro has a built-in 

pendulosity of 5 x 10~2 N-m/g. The outermost cylindrical housing is approximately 3V4 in. (82 mm) long by 

nearly 2V2 in. (62 mm) in diameter. The overall unit 

weighs 1.8 Ib (0.82 kg), and its life expectancy, under

continuous operation, is more than 5 years (44,000 

hours). The design objective on sensitivity is to hold it 

to a small fraction of one-millionth of the force due to 

gravity.

In the mechanical design of the IMU, the three accelerometers are to be arrayed as shown schematically 

in the lower part of figure 6. This is an obvious departure from the more conventional arrangement in which 

the three sensitive, or input, axes of the accelerometers 

would have been oriented along the three basic Cartesian coordinate reference axes. In that conventional 

arrangement, the input axes of two accelerometers 

would be alined with the north- and east-pointing axes

INPUT AXIS

RESOLVER SLIP RING 

ASSEMBLY

INPUT AXIS 

DRIVE MOTOR

PIG MOUNTING 

DISC

PENDULOUS 

INTEGRATING 

GYRO ASSEMBLY 

(PIG)

FLANGE 

MOUNT

DUPLEX 

BEARING

OUTER 

CASE

SHROUD

i

* ^i T + f?

FIGURE 9. The pendulous integrating gyro accelerometer.
INSTRUMENT SYSTEM 19

of the IMU and, thus, would lie in the horizontal plane 

established by the stable platform. This would mean 

that the corresponding orthogonal unbalanced spin 

axes of the two accelerometers would sense no acceleration due to gravity, and, so, during much of the total 

operating period for the IMU, the respective two 

output-axis shafts would rotate very little. Yet for the 

accelerometer, the best performance is obtained when 

the output shaft experiences some rotational movement enough to overcome the static friction of the 

bearing. Hence, the deliberate choice to orient the triad 

of accelerometers so that, in addition to the required 

orthogonality of the three input axes, each would continuously sense an equal share of the local gravity vector; that is, 0.58 g. If the origin of the three orthogonal 

reference axes is imagined as the corner of a cube, then, 

by placing in a vertical position the longest cube diagonal from that corner point, the correct orientational 

geometry is achieved. This means that each accelerometer input axis is at an angle of 54°44' with the local 

vertical (gravity vector) or, as shown in figure 6, at an 

angle of 35°16' with the horizontal plane of the stable 

platform.

The accelerometer triad occupies its own physically 

separate section of the stable platform and, by adopting 

a conical shape for this section, the thermal paths from 

accelerometers to enveloping heat exchanger (see discussion under "Temperature Control") are made short. 

For the same reason, shortness of thermal paths, the 

three required electronics modules are mounted between the accelerometer units; this also serves to minimize lengths of electrical leads.

Gimbal System

In the discussion on gyros, the frame that supports a 

gyro was identified as a gimbal. In the IMU, the function of the gimbals transcends that of simple support 

and extends to the more basic purpose of decoupling or 

isolating the gyros from the angular motion of the overall enclosing case mounted on some rigid base. For the 

APT system, the base is to be shock mounted on the 

cabin floor of a relatively light twin-engine aircraft. 

Nothing in the intended flight missions will require 

any violent maneuvers of the aircraft. Thus, the IMU 

gimbal structure can be simply that which is needed to 

isolate the stable platform, instrumented by the three 

gyros, from the standard operating three-dimensional 

angular motion (roll, pitch, and yaw) of the aircraft. As 

a matter of fact, slip rings provide unlimited freedom 

for the aircraft to yaw (turn), and freedom of plus or 

minus 190° is provided for it to roll and pitch.

The foregoing description of the operating environment for the APT system led to the decision to isolate 

the stable platform with a three-axis gimbal structure,

the arrangement of which is shown schematically in 

figure 10. Starting at the inside and proceeding outward, the three axes in sequence are the vertical or 

azimuth axis (color coded in black), the elevation axis 

(color coded in red), and the roll axis (color coded in 

blue); the termini of these axes are seen to be in the 

elevation gimbal, the roll gimbal, and the enclosing 

case (also color coded in black), respectively. The 

arrowheads denote the positive-direction conventions. 

The two ends of each axis fit into a pair of preloaded 

ball bearings which, in turn, are mounted in steel 

sleeves to minimize change in preload with temperature. As an added precaution against change in preload 

through temperature differentials between gimbals, 

each gimbal is designed with a diaphragm at one end.

Mounted in the appropriate gimbal, or case, at one 

end of each of the three axes is a torque motor that 

provides a driving torque around the axis (fig. 10). In 

the APT system, the motor for the azimuth axis has a 

torque capability of 200 oz-in. (1.41 N-m), and the motors for the elevation and roll axes each have torque 

capabilities of 30 and 60 oz-in. (or 0.21 and 0.42 N-m), 

respectively. Each motor has a friction torque of less 

than 1.5 oz-in. (0.01 N-m). Modules containing the 

torque motors are shielded magnetically from other 

components in the system.

Mounted in the appropriate gimbal at the opposite 

end (from the torque motor) of each of the three axes is 

a precision multispeed resolver (fig. 10) to measure the 

angular position of the gimbal or case referenced to the 

three-coordinate-axis system (x, y, and z) being carried 

by the stable platform.

The gyro section, the accelerometer section, and 

their required electronics modules are all assembled, 

as part of the stable platform, into the elevation gimbal. The conical shape of the accelerometer section provides added stiffness for the stable platform. For 

maintenance service to the stable platform, access may 

be gained simply by separating the gyro and accelerometer sections. Gimbal and stable platform material is 

"6061" aluminum, which was chosen for its low density, good thermal conductivity, high mechanical stability attainable through heat treatment, and high resistance to corrosion. The natural resonant frequency 

of each gimbal is in the range of 75 to 100 Hz. Total 

weight of the stable platform is 36 Ib (16.3 kg).

Resolvers

In the APT system, the sophistication of the IMU as 

a high-precision measuring device is enhanced through 

the ability of resolver components in the gimbal support structure to sense and read out with extreme accuracy the angles linking orientation (attitude) of the 

carrying vehicle (aircraft) with reference to the x, y,
20 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

and z coordinate axes of the stable platform. This angle 

information is vital to the operation of the laser tracker 

component in the APT system, which is described in 

the section titled "Laser Tracker."

Among electrical devices, resolvers are really variable transformers that are descendants of earlier 

(World War II) computing units in gun directors that 

generated pointing data from radar moving-target deIMU ELEVATION 

GIMBAL 

AZIMUTH AXIS 

TORQUE MOTOR

ELEVATION AXIS 

TORQUE MOTOR

ELEVATION OR 

MIDDLE AXIS

DIRECTION 

OF FLIGHT

ROLL OR 

OUTER AXIS

ROLL AXIS 

RESOLVER

ROLL AXIS 

TORQUE MOTOR

STABLE 

PLATFORM

AZIMUTH AXIS 

RESOLVER

ELEVATION AXIS 

RESOLVER

VERTICAL (AZIMUTH) 

OR INNER AXIS

SUPPORT 

HOUSING 

OR CASE

SUPPORT HOUSING 

OR CASE

AIRCRAFT CABIN FLOOR

FIGURE 10. Support structure for inertial measurement unit.
INSTRUMENT SYSTEM 21

tection equipment. In essence, each resolver in the APT 

system comprises two concentric rings; the inner is 

called the rotor, and the outer, the stator. The rotor is 

mounted on a movable axis of the gimbal-support structure, and the stator is mounted in the companion gimbal where that particular axis terminates. The term 

"resolver" originated from its function of resolving 

electrical signals into components in accordance with 

the angular rotational movement of rotor relative to 

stator.

In a fundamental and elemental way, the rotor and 

stator may be visualized as carrying two distinct and 

separate coil windings apiece. The windings in a given 

pair are at right angles to each other as shown in the 

schematic of figure 11. In this schematic, the rotor and 

stator are drawn separately and side by side, but, in the 

actual resolver, they are concentric. If the rotor is 

imagined as the primary of an electrical transformer 

and its pair of windings energized with alternating 

electrical currents (at voltages ei and e2 in fig. 11), then 

two magnetic fields are produced at right angles to 

each other. Field strength is related directly to number 

of turns in the winding and amperage of the impressed 

electrical current. These two magnetic fields of the 

rotor induce two corresponding output voltages (e0l and 

e02 in fig. 11) in the pair of windings carried by the 

stator which can now be imagined as fulfilling the role 

of the transformer secondary. The highest voltage a 

primary winding can induce in a secondary winding is

SECONDARY

COIL

WINDING

O

T602

1

e2

cos Gsin 0 0)

WHERE

k= ELECTRICAL COUPLING COEFFICIENT 

BETWEEN ROTOR AND STATOR

 = "IN PHASE" 

FIGURE 11. Resolver schematic and nomenclature.

when the two are parallel to each other; the induced 

strength dwindles to zero when the two orientations 

are mutually perpendicular. Thus, with particular reference to figure 11, if the rotor turns through an angle, 

0, with respect to the stator, then the induced (output) 

voltages, e0l and e02, may be computed precisely from 

the equations shown.

To meet the overall performance specifications for 

the APT system, 0 must be related to the foregoing 

determinations of voltage to a precision of 4 arcseconds. To comprehend the real significance of this 

precision visualize a circle of 1-in. (25-mm) radius. A 

1-in. chord will obviously subtend a central angle of 60° 

and, through simple proportion, a central angle of 

4 arc-seconds, therefore, represents a chord length of 

only 18-millionths of an inch (4.6 x 10~4 mm). For the 

APT system, the resolver radius, given as a convenient 

round number, is 2 in. (51 mm) and, thus, an angular 

displacement, 0, of 4 arc-seconds signifies a chord 

length of only 36-millionths of an inch (9.1 x 10~4 mm).

As a way of ensuring that the indicated level of precision is reached, the angular movement of the rotor is 

multiplied electromechanically in a way somewhat 

analogous to use of a suitable gear train. Pursuing that 

analogy, a gear ratio of, for example, 36 to 1 can be 

effectively obtained by adding a second set of windings 

in the design of the resolver such that a full 360° of 

electrical rotation, and, hence, signal resolution, takes 

place every 10° of rotor rotation. The result is described 

as a single-speed coarse and 36-speed fine output capable of providing the required accuracy for the APT 

system.

The 4-arc-second angular displacement referred to 

above is actually sensed as a rate of change in magnetic 

flux in the 36-speed resolver windings, which is observed as a proportional change in voltage of about 0.5 

millivolt (mv) at the terminals of those windings. Electronic amplifiers are readily capable of precisely detecting and faithfully processing and relaying signals 

originating at these fractional millivolt levels.

Added insurance for maintaining high precision is 

gained by mounting rotor and stator in a steel ring to 

minimize thermal stress in their magnetic iron parts. 

Thermal tests have shown that the ambient temperature can range from 0°F to 120°F (-18°C to 49°C) with 

an error of less than 3 arc-seconds appearing in the 

resolver readout. Magnetic shielding also is designed 

around each resolver to eliminate or minimize errors 

caused by external magnetic fields. Finally, the structure and geometry of each resolver are shaped very 

precisely and are virtually distortion free. Similarly, 

the electromagnetic circuits are built very precisely for 

low power loss, symmetry, and stability. Principal features of a resolver are as shown in figure 12.
22 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

o 

ex 

ex

Q
INSTRUMENT SYSTEM 23

Feedback Loops

In preceding discussions, particularly those concerning resolvers, accelerometers, and gyros, references 

have been made to electromechanical elements whose 

function it is to measure angular displacements very 

precisely. Little has been said, however, as to the disposition of such information. An important ingredient in 

the ability of a stable platform to remain stable is not 

only a capability for sensing the magnitude and direction of any tilt (angular deviation) away from the 

desired stable orientation in three-dimensional space 

but also a capability for followthrough with appropriate corrective action. The followthrough is accomplished by way of servoamplifiers and drive motors that 

deliver electrical signals and mechanical torques or, as 

the entire process may be labeled, by way of feedback 

loops.

Basically, a feedback loop is a form of automatic control intended not only to sense the difference between 

an input, or reference, signal and an output, or status, 

signal but also to reduce this difference, or deviation, as 

it may be called, to near zero or acceptable limits. In 

examining the simplified electromechanical feedback 

loop, shown schematically in figure 13A, start at the 

electrical input signal, Vj, which is passed through the 

deviation sensor, a device that algebraically sums or 

compares it with the electrical signal, Vf, being fed 

back through the loop. Now, proceeding clockwise 

around the loop, the resultant deviation signal, V,j,

MECHANICAL OUTPUT

FEEDBACK SIGNAL LOAD MOVEMENT DATA

MECHANICAL OUTPUT 

ELECTRICAL SIGNAL

ELECTROMECHANICAL FEEDBACK LOOP

MECHANICAL OUTPUT

DEVIATION SENSOR

FEEDBACK SIGNAL

B

GYRO FEEDBACK LOOP

TILT REMOVAL DATA 

 MECHANICAL OUTPUT 

= ELECTRICAL SIGNAL

FIGURE 13. Schematized feedback loops.

from the deviation sensor is fed electrically into a servoamplifier from which suitable energy flows to a 

motor which must provide the mechanical muscle output to operate a controlled element in some desired 

way. As this element is moved, an electrical pickoff 

device generates load movement data which, in effect, 

is the feedback signal, Vf, which is relayed to the deviation sensor, thereby closing the loop. Thus, the forward part of the loop performs the control function, and 

the back part,the status-report function. The loop system works continuously to bring the deviation toward 

zero; that is, to reach and maintain a null position or 

balance so that Vj Vf. This action of the loop not only 

overcomes internal forces, such as friction, but also opposes external disturbances, all of which tend to force 

the controlled element into undesired movements or 

positions.

Two characteristics are significant in describing 

feedback loops. First, the controlling signal is not only 

a measure of a deviation, but it triggers a response in 

the direction needed to diminish that deviation. Second, design art must be applied to a determination of 

the rate at which the deviation is to be diminished so 

that the controlled element or load does not overshoot 

the desired position. Commonly, the disturbances acting on the controlled element are changing continuously, and, ideally, the feedback loop will follow those 

changes closely. However, as the loop response is quickened, the entire system bcomes less stable, and a point 

is reached at which it can be made to break into oscillation, or jitter. Thus, the real art is in designing the loop 

to be fast enough to perform its assigned task while 

properly limiting overshoot so as to remain stable. In 

general, stability is assured by incorporating compensation elements which precisely set the phase and amplitude relations throughout the loop.

In the IMU for the APT system, feedback loops play 

a vital role in enabling the stable platform to maintain 

continuously its proper orientation in threedimensional space. A convenient example of one such 

loop may be taken arbitrarily from the stabilization 

servo loop designed and constructed to serve one of the 

three gyros on the stable platform. In this case, the 

desired position is given by the gyro null (Vi = 0). Any 

incipient tilt of the stable platform in any particular 

direction may be subdivided into three orthogonal component tilts (rotational movements) around the three 

orthogonal input (sensitive) axes of the gyros. Now, 

consider just one given gyro (fig. 13B ). It will detect the 

incipient or small component tilt around its input axis 

and feed this information, as an angular deviation signal, Vf, back to the deviation sensor. Because Vj is zero, 

the output of the deviation sensor is simply (-Vf), the 

deviation signal, which is fed to the servoamplifier. 

Continuing clockwise around the schematic (fig. 13B),
24 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

the amplifier feeds electrically into a gimbal torque 

motor which provides the amount of power needed to 

rotate the platform and to begin to offset the component 

of tilt the gyro has detected. The electrical pickoff in 

the gyro monitors the restorative movement and relays 

these data on around the loop as a new and diminished 

feedback signal, Vf, into and through the deviation sensor for repetition of the cycle, if necessary. The system 

works continously to eliminate any angular deviation 

and to reach and maintain the desired null position. In 

doing this, the loop must oppose the variety of extraneous factors, such as internal friction or air turbulence 

on or motion of the carrying aircraft, that materialize 

as disturbance torques on the stable platform, tending 

to force it into undesired movements or positions.

Although figure 13B may be oversimplified, the principles are valid, and, if three such individual feedback 

loops are envisioned to serve the three respective gyros, 

then their combined performance provides the capability for holding the stable platform continuously in the 

desired orientation. This design is more complicated, 

however, than immediately meets the eye. Because the 

three loops act on a common element, the stable platform, and,because some nonorthogonality may have 

been built inadvertently into the platform mechanization, the loops could interact in the sense of cross coupling, which could then lead to oscillatory instabilities. 

This does not happen, however, simply because the 

problem has been recognized, studied, and solved by 

matching the dynamic characteristics of the respective 

loops against each other. Thus, the design choices have 

been made from a broad background of experimental 

data and experience.

Electronics

In the IMU, the electronic design concepts emphasize 

nonlinear techniques to minimize electrical power 

losses and standardization (insofar as possible) through 

diligence in recognizing common requirements. The 

electrical circuits are partitioned so that only the precision control portions associated with the inertial instruments are located on the stable platform. This includes circuits for a telemetry system to provide 

state-of-health information about the inertial components. Other circuits, such as those involving power 

supplies and resolver encoders, are kept external to the 

stable platform and gimbal system. The external circuits also include the capability for torquing the gyros.

The electronic design embraces electrical circuits 

that are the means for creating the stabilization subsystem, which must isolate the stable platform from 

aircraft angular motion as has been described functionally in the preceding section. Three individual stabilization circuits or loops (azimuth, elevation, and roll)

are needed to separate out all possible angular motion 

of the aircraft and to maintain the stable platform at 

the desired orientation. The three individual loops are 

shown schematically in figure 14, but their explanation should be followed with frequent reference to the 

schematic for the IMU support structure (see fig. 10).

Note first (fig. 14) that the deviation signals for instructing the gimbal torque motors come from the gyro 

signal preamplifiers located on the stable platform. 

Now, with respect to the azimuth stabilization loop, 

obviously any "yaw" motion of the aircraft simply 

tends, because of slight amounts of friction in the gimbals, to drag the stable platform around its azimuth 

axis, and the deviation signal from the z gyro is, therefore, fed (fig. 14) directly around the loop to the azimuth axis torque motor to null out the incipient gyro 

motion.

Next consider the elevation stabilization loop which 

obviously concerns aircraft "pitching" motion. Such 

motion tends, because of friction, to drag the elevation 

gimbal around the elevation axis which can prompt 

deviation signals from either one or both of the x and y 

gyros (depending on aircraft heading) as the stable 

platform begins to tilt away from its locally level position. These signals (proportional to the x and y components of the pitching motion) are fed to the primary 

windings of the gyro signal resolver (fig. 14). One secondary winding is alined with the elevation axis to 

supply the elevation component signal that traverses 

the elevation loop to the elevation axis torque motor. 

The object is to null out the incipient pitching motion 

as measured by either one or both of the x and y gyros 

as appropriate.

Finally, consider the roll stabilization loop which obviously concerns aircraft "rolling" motion. Such motions tends to drag the roll gimbal around the roll axis 

which, again, can prompt deviation signals from either 

one or both of the x and y gyros (depending on aircraft 

heading) as the stable platform beings to tilt off level. 

The roll signal is supplied by the other secondary winding of the gyro signal resolver and traverses the roll 

loop to the roll torque motor. And, again, the object is 

to null out the incipient roll motion as measured by 

either one or both of the x and y gyros as appropriate. 

The reduction in gain in the roll loop, due to aircraft 

pitch angle, has no practical effect on servo performance in isolating the stable platform from aircraft 

motion.

Just before the start of a flight mission, the IMU 

must be leveled approximately and oriented with respect to geographic north and east in preparation for its 

self-alinement procedures. Thus, provision is made for 

operating the three gimbal servoamplifiers and their 

torque motors independently in an angle-command 

mode under computer control. In this mode, the error 

signal for each servoamplifier is formed when the com-
INSTRUMENT SYSTEM 25

puter determines the difference between the angular 

position input from each two-speed resolver on the 

three gimbal axes (see fig. 10) and the respective command angles that the APT system operator has entered 

by way of the keyboard. The three error signals then 

are transmitted to the appropriate servoamplifiers 

through the digital-to-analog (D/A) output channels 

(see "Computer" section), and, thus, each gimbal is 

driven to the corrected commanded angle. Command 

angles also may be generated by a computer program 

that allows for automatic operation. For testing purposes without the computer, the gimbal angles can be 

commanded manually using hand-settable resolvers as 

control transformers to generate the error signals in a 

fully analog configuration.

Temperature Control

Work began very early in this project to shape thermal design concepts that would assure control of the 

stable platform temperature environment to a high degree of precision. Alinement stability of the stable platform is a function of gyro and accelerometer performSTABLE PLATFORM

ance, which, in turn, is extremely sensitive to 

temperature change. The concepts that evolved resulted in the thermal control system shown schematically in figure 15.

Immediately evident is the fact that a high operating 

temperature in the compartment occupied by the stable 

platform (fig. 15) allows aircraft cabin air to be the 

ultimate heat sink over a wide range of ambient temperatures [up to, for example, 115°F (46°C)]. Recall that 

the gyro and accelerometer rotors are operated in floats 

suspended in a viscous fluid held at a constant temperature of 139°F (59°C). The progressive downward steps 

in environmental temperature, proceeding outward 

from rotors to aircraft cabin (fig. 15), are as labeled or 

can be approximated. Strategic placement of small 

electrical heaters along the thermal gradient guarantees the needed precision in temperature control regardless of gimbal orientation. Thus, for much of the 

expected range of field-operating conditions, heat will 

be supplied to the instrument system, and the need for 

a cooling system is obviated.

The inertial instruments (gyros and accelerometers) 

and their internal electronics are mounted on the staELECTRONICS

ELEVATION 

LOOP

ROLL 

LOOP

AZIMUTH 

LOOP

EXPLANATION

TM= TORQUE MOTOR

GSR = GYRO SIGNAL RESOLVER

FIGURE 14. Electrical circuits for stabilization subsystem of inertial measurement unit.
26 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

ble platform in a way that permits all the heat they 

generate to be conducted to the stable platform heat 

exchanger, which completely surrounds and is integral 

with the platform. High-wattage dissipators in the 

electronics package are placed closest to the heat exchanger to shorten thermal paths and to minimize 

thermal disturbances. Voids in the entire enclosure are 

filled with insulation to suppress convective air currents around the platform.

A constant-speed fan, mounted on the elevation gimbal, provides steady air flow at a constant temperature 

through the stable platform heat exchanger (fig. 15).

As it transits the exchanger, the air collects and carries 

away heat from the platform The air then passes 

through the inner chamber of the heat exchanger that 

is wrapped around the elevation gimbal, losing heat by 

conduction to the outer chamber of the exchanger during this passage. This inner air loop is completed when 

the cooled air moves back to the intake side of the 

constant-speed fan to repeat the cycle.

The rate of movement through the outer air loop will 

vary, depending on the temperature differential between the outer chamber of the heat exchanger on the 

elevation gimbal and the ambient air in the aircraft

VARIABLE SPEED FAN

AZIMUTH AXIS 

AIR OUTLET

SUPPORT HOUSING 

OR CASE

HEATER AND 

THERMAL INSULATION

INLET 

AIR FILTER

HEAT EXCHANGER 

OUTER CHAMBER^

STABLE PLATFORM 

HEAT EXCHANGER

HEAT EXCHANGER 

INNER CHAMBER

VARIABLE FAN SENSOR

INLET 

AIR FILTER

CONSTANT SPEED FAN]

FIGURE 15. Thermal control schematic for inertial measurement unit
INSTRUMENT SYSTEM 27

cabin. Thus, a variable-speed fan, also mounted on the 

elevation gimbal, is required (fig. 15) to circulate the 

air in this outer loop. Suitable control of the variable 

speed is achieved through a thermistor sensor placed in 

the inner air loop. By choosing its location (fig. 15) in 

the inner chamber of the dual-chambered heat exchanger, it serves as an early-warning device to forestall any change in air temperature that might alter 

the steady and known rate of heat dissipation required 

for the stable-platform enclosure.

Evolution of the preceding thermal control concepts 

had considerable impact on the design shapes for the 

stable platform and the internal electronic modules 

that are mounted thereon. Figure 15 is highly schematic to illustrate the thermal concepts. A more detailed and realistic view is shown in figure 16, which 

also identifies the inner and outer air loops.

In brief, air moves around the inner loop from the 

output of the constant-speed fan (at top center of the 

elevation gimbal) downward through the inner chamber of the dual-chambered heat exchanger, then upward through the heat exchanger that encloses the stable platform, and finally back to the intake side of the 

fan.

Air moves around the outer loop from the output of 

variable-speed fans (near the top of the elevation gimbal) downward through the enclosure formed by the 

support housing, or case, then upward through the 

outer chamber of the dual-chambered heat exchanger, 

and finally back to the intake sides of the fans. Although two variable-speed fans are shown in the vertical section of figure 16, some tests have suggested that 

a three-fan array may be preferable, spaced 120° apart 

in plan view, for better symmetry of air flow. When a 

prototype of the APT system is built entirely to the 

basic design specifications, it will feature the most appropriate array.

In the realm of temperature control, the outer air 

loop is really a coarsely controlled buffer zone between 

the outside world (aircraft cabin), where the virtually 

uncontrolled ambient temperature range is large, and 

the inner air loop, where control is precise and tight. 

The outer air loop connects with cabin air through an 

annular plenum chamber, which is protected by the 

annular inlet air filter shown in the lower part of figure 

16. The fan shown passing through this chamber and 

filter can exhaust air to the cabin from the outer air 

loop. The coarse control, whereby the temperature 

range is held between 110°F (43°C) and 120°F (49°C), is 

simply the manipulation, as needed, of this fan and 

electrical heaters in the outer loop.

The foregoing details concerning temperature control constitute a fairly rigorous portrayal of what was 

envisioned in the original design for the APT system. 

The changes wrought by the decision to build the first 

prototype with surplus Air Force gyros and accelerometers are summarized most significantly by noting the 

following temperature regimen that is to be maintained:

1. Viscous fluid which buoys floats for gyro and accelerometer rotors, now to be held constant at 

135°F (57°C).

2. Inside stable-member package (see figs. 15,16), now 

to be held at 180°F (42°C).

3. Outer air loop (figs. 15, 16), which is the space between the elevation gimbal and the support housing, or case, now to be held at or near 66°F (19°C). 

Cooling coils, built into the housing, will circulate 

fluid at 59°F (15°C), as needed.

4. Aircraft cabin air, around the outside of the housing, will now be conditioned (heated or cooled), to 

be held in the range 70°F (21°C) to 80°F (27°C).

Performance Dynamics

The preceding sections have described the principal 

individual components, structures, and subsystems 

that comprise the IMU, the heart of the inertial navigator. In the functioning airborne APT system, the inertial navigator must be supplemented by other elements, such as the laser tracker and the onboard data 

processing unit. Before venturing into descriptions of 

these other key elements, however, it is appropriate to 

introduce and explain some of the performance dynamics that will be required of the inertial navigator. The 

intent is to explain by relying primarily on physical 

and geometric concepts that appeal to the intuition 

rather than on rigorous mathematical proofs. The latter are available in a number of excellent textbooks, 

several of which are in the list of references at the end 

of this paper (see especially Britting, 1971; Broxmeyer, 

1964; Pitman, 1962).

A variety and substantial number of inertial navigation systems have been conceived, designed, and built 

over the past three decades. Most system types have 

been demonstrated successfully in flight, and some are 

now in regular production and use. The choice of a 

system depends on the particular flight application intended, and, as stated earlier (see section titled 

"Inertial Navigator"), the IMU for the APT system was 

chosen to be mechanized in a tangent-plane orientation. This means that, to a very good approximation, 

two axes (x and y) out of the set of three orthogonal 

axes, created by the stable platform in the IMU, will at 

all times define a plane perpendicular to the actual 

gravity direction at the position where the alinement 

took place.

All inertial systems require measurement of the 

magnitude and direction (referenced to the stableplatform x-y-z axis orientation) of the specific force vector acting on the carrying vehicle, and all must cope 

with the fact that the force of gravity is indistinguish-
28 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

able from the force associated with acceleration of the 

vehicle when it is in motion; that is, only their vector 

difference can be measured.

Because the inertial navigator in the APT system 

will always operate in a near-Earth environment, with 

emphasis on high-precision measurements in a relatively localized survey area, it is especially important 

to develop first a clear understanding of how the inertial navigator "feels" or senses that environment before

SUPPORT HOUSING 

OR CASE

VARIABLE-SPEED FAN

AIR LOOP, INNER 

AIR FLOW

any motion of the carrying vehicle is considered. This is 

tantamount to defining the initial or starting conditions for the inertial navigator because it is basically a 

dead-reckoning device, and these initial data must be 

given to at least the precision required during subsequent operation of the system as a navigator.

Every point on Earth has a gravity vector associated 

with it; that is, the acceleration due to gravity has a 

specific magnitude and direction. When the vehicle

AZIMUTH AXIS

CONSTANT-SPEED FAN

VARIABLE-SPEED FAN 

AIR LOOP, INNER

AIR LOOP, OUTER

DUAL-CHAMBERED 

HEAT EXCHANGER

128°F±0.1°F 

(53 °C)

ELEVATION 

GIMBAL

ROLL 

GIMBAL

INLET AIR 

FILTER

ELECTRONICS 

PACKAGE

LASER

TRACKER

ASSEMBLY

ELEVATION AXIS

AIR TEMPERATURE 

110°Fto 120°F 

(43 °C) (49 °C)

FAN

FIGURE 16. Support and thermal control structure for inertial measurement unit.
INSTRUMENT SYSTEM 29

carrying the inertial navigator is not in motion, the 

accelerometers will sense and measure only the gravity 

vector, and, in fact, this information is used initially to 

level the stable platform. In a tangent-plane orientation, the x y plane of the stable platform senses no 

component of acceleration due to gravity. In the APT 

system, because of the geometry of the chosen accelerometer array (see fig. 6), the stable platform is 

tangent-plane level when the three accelerometers 

sense equal components of the specific force due to 

gravity, and the z axis, thereby, is alined with the actual gravity vector.

The rate of the Earth's rotation about its polar axis is 

a known quantity that is everywhere the same, and the 

motion of any point on the surface is easterly. Thus, it 

can be said that every point on Earth has the same 

angular rotation vector associated with it.

The aforementioned two vectors, gravity and Earth 

rotation, dominate in characterizing the dynamic 

Earth environment. Further discussion is needed, however, to explain more adequately how the inertial navigator is affected by and senses them and how the 

stable platform must be commanded to maintain continuously the desired orientation x axis pointing 

north, y axis pointing east, and z axis pointing downward along the local vertical at the alinement point 

regardless of the local area where the instrument is 

operated on or above the Earth. If it were not so commanded, then its gyros would know nothing better than 

to hold the x-y-z axis orientation of the stable platform 

inertially fixed in whatever startup orientation was 

given, despite the steady and continuing rotation of the 

Earth. If, in that situation, the vehicle carrying the 

inertial navigator stood still for 12 hours (one-half a 

revolution of the Earth), this would provoke the 

graphic discovery that the stable platform was by then 

virtually upside down it would not have changed its 

orientation in space to keep up with the rotating Earth. 

As the following discussion will reveal, all the information needed for the initial self-alinement of the x, y, and 

z axes is contained in the accelerometer signals. The 

stable platform is simply turned to and held in the 

correct orientation by applying, around the appropriate 

axes, gyro torques, the magnitudes of which are calculated by the onboard computer in accord with the functional relations that are based on the accelerometer 

signals.

In seeking further enlightenment, through the following discussion, reference to a cutaway perspective 

view of an idealized and homogeneous ellipsoidal Earth 

(fig. 17) is useful. Note first how the stable-platform set 

of three orthogonal axes (in dark blue) must be drawn 

at each of the several points P1? P2, and P3, shown in 

figure 17, to maintain correctly the stipulated orientation. For convenience, P1? P2, and P3 are chosen on the 

Earth's surface in the northern hemisphere.

Several significant facts stand out in figure 17; for 

example, observe how the direction of the local vertical, 

the z axis, which is normal to the ellipsoid surface, 

changes with the location of P1} P2, and P3 . As a matter 

of fact, the definition of astronomical latitude, of special importance in geodetic surveying, is the angle, 3>, 

that the local vertical (z axis) makes with the equatorial plane. Thus, <£ ranges between the limits of 0° at 

the equator, where the z axis lies in the equatorial 

plane (see fig. 17), to 90°N at the north geographical 

pole, where the z axis is coincident with the polar axis 

or perpendicular to the equatorial plane. Note also 

that, in the definition of latitude, the local vertical need 

not, and, in the real world usually does not, pass 

through the Earth's center of mass, C, except at the 

equator and poles.

Observe, also (fig. 17), that P1? P2, and P3 , are chosen 

along a single meridian at locations, in the same respective order, at the equator, at an arbitrary middle 

latitude, and at a short distance from the north geographical pole. If the set of three orthogonal axes is 

studied at each point, then the difference in orientation 

between any two points is seen to be described as a 

rotation of the set around the y axis through an angle 

equal to the change in latitude between those points. 

Here, then, is a fundamental piece of orientation 

knowledge for the set of three axes in assessing the 

effect of "northward" or "southward" components of 

changes in location of the inertial navigator on or over 

the Earth's surface.

Now, if a similar reasoning process can be developed 

for the eastward or westward components of change in 

inertial navigator location, then the general base will 

have been completed for understanding any conceivable change in location. Immediately, however, eastward and westward components are seen to connote 

some sequence of points P1? P2, and P3 on a single parallel of latitude. The reasoning then becomes complicated by the subtleties of the convergence of the meridians as successive northerly parallels of latitude are 

involved. The situation can be neatly bracketed, however, by describing two limiting cases.

First, consider P1} P2, and P3 as lying on the equator 

(<j> = o°). Imagine further that, at each point, a set of 

three axes (x, y, and z) is drawn in the previously specified orientation. Now it can be seen (fig. 17) that, as the 

orientations at P1? P2, and P3 are examined, the z axis 

at each location lies in the equatorial plane and is perpendicular to the equator (locally vertical). For the z 

axis to maintain this perpendicularity, with a change 

in location along the equator, requires simply a rotation of the set of axes "around the x axis" an amount 

equal to the angular change in longitude.
30 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

Second, consider the case where P1? P2 , and P3 are 

lying on a parallel of latitude just below the north geographical pole, near a latitude, for example, between 

85°N and 90°N (85°N «D< 90°N). Again, imagine a 

set of three axes drawn at each point. Now, it can be 

seen that, as the orientation of the set is examined 

successively at P1} P2 , and P3 , the z axis moves very 

little to remain locally vertical; the principal rotation 

for the set of axes must be "around the z axis" so that 

the x axis can be kept pointing toward the north pole. 

In the limit, as the latitude approaches 90°N, this angular amount would equal the angular change in longitude, and, at the pole, the convergence of the meridians 

equals the angular change in longitude.

For any parallel of latitude intermediate between 

the two foregoing limits, as the orientation is examined 

at P1? P2 , and P3 , the required rotation of the set of 

three axes will involve some combination of two angular components around just the x and z axes. Quantification of these two particular components for the respective two axes now merits interpretation.

With no change in latitude involved in east-west 

changes in location, the z axis maintains a constant 

angle with the equatorial plane (fig. 17). This means 

that no rotation is needed around the y axis and affirms 

the statement in the preceding paragraph that, if the 

set of three orthogonal axes is studied for any two 

points, P, at the same latitude, then the difference in

EARTH'S POLAR AXIS

DIRECTION OF 

EARTH'S ROTATION

LOCAL VERTICAL 

(GRAVITY VECTOR) 

AT POINT P3 

(LAT.=03 >

LOCAL VERTICAL 

(GRAVITY VECTOR) 

AT POINT ?

ELLIPSOIDAL EARTH

MERIDIAN 

-EARTH'S CENTER

EQUATOR

EQUATORIAL 

PLANE

LOCAL VERTICAL 

(GRAVITY VECTOR)

AT POINT P.

(LAT.= 0°)

FIGURE 17. Orientation of locally level three-coordinate reference scheme at selected points on an idealized ellipsoidal Earth.
INSTRUMENT SYSTEM 31

orientation is simply accounted for as a rotation of the 

set by some angular amounts around the x and z axes. 

The angular amount around the z axis (Rotz) can be 

equated immediately to the convergence, 6, of the meridians between the two points. This follows by visualizing the spherical geometry in figure 17 and by noting 

that, with no rotation allowed around the y axis, the 

only rotation that can keep the x axis pointing north is 

that which occurs around the z axis, and it must equal 

6. Geodesy texts (see, for example, Hosmer, 1948, p. 

288) give the value of G (for a spherical Earth) as the 

product of the angular change in longitude, AX, and the 

sine of the chosen parallel of latitude. In simple mathematical form, this means

Rotz = 6 = |AX sin $| .

This relation can be tested immediately for agreement 

with the descriptions already given on how the set of 

three coordinate axes must be rotated for east-west 

changes in location of the inertial navigator along either of the two limiting parallels of latitude. Thus, at 

the equator (4> = 0°) obviously 6 will be zero, which 

signifies no rotation around the z axis as alternate locations are considered around that latitude. At the other 

limit, as 4> approaches 90°N, 6 closely approaches AX, 

which signifies that, in the limit, all the rotation will 

be around the z axis. Agreement, therefore is achieved. 

All that remains now is to specify, in the above situation, the angular amount of the required rotational 

component about the x axis (Rotx). Again, it is important to visualize the spherical geometry by studying 

figure 17. Note that Rotx is the only means (with no 

rotation allowed around the y axis) for keeping the z 

axis pointing in the direction of the local vertical, and 

the magnitude of this rotation is measured by the 

length of circular arc, along the parallel of latitude, 

that represents the eastward or westward change in 

location from one point to another. This circular arc is 

the measure of the angle through which the z axis must 

be swept to stay properly oriented as a consequence of 

the eastward or westward change in location. For a 

given change in longitude, AX, the length of circular arc 

is obviously a function of the latitude, 4>, and, from 

texts on spherical geometry (see, for example, Ayres, 

1954, p. 188), this is given (for a spherical Earth) as the 

product of AX and the cosine of the particular parallel 

of latitude. Thus, in simple mathematical terms,

Rotx = |AX cos O| .

If this relation now is tested for agreement with the 

earlier observations on conditions to be met for eastwest differences in location of the inertial navigator 

along either of the two limiting parallels of latitude,

then there follows:

1. At the equator (4> = 0°), Rotx = |AX|.

2. Near the pole ($^90°N), Rotx ^ 0. 

Once again, agreement is achieved, and the foregoing 

developments have supplied the needed corollary piece 

of orientation knowledge for the set of three axes, in 

assessing the effect of "eastward" or "westward" components of changes in the location of the inertial navigator on or over the Earth's surface.

In a summary and generalized way, therefore, the 

change in orientation of the set of three orthogonal 

axes, with changes in location of the inertial navigator, 

is described in the following succinct form: For northward (+) or southward ( ) components of location 

change,

Rotx = 0, 

Roty = -AO, 

Rotz = 0.

For eastward (+) or westward (-) components of location change,

Rotx = AX cos <£,

Roty = 0,

Rotz = -AX sin <£.

Note the positive and negative conventions adopted for 

changes in location in each of the four cardinal direction. As for angular rotational movement of the set of 

three axes, the positive sense is in accord with the 

right-hand-thread rule. Heretofore in this section, only 

rotational magnitudes have been shown with no regard 

for positive or negative directions.

To a degree, the preceding discussion has anticipated 

how the inertial navigator senses the Earth's rotation. 

Obviously, if eastward changes in location are analyzed with respect to how the orientation of the x-y-z set 

of axes in the stable platform must be changed, then 

this is very close to analyzing for Earth rotation. The 

picture can be enhanced, however, by observing 

(fig. 17) how the Earth's rotation is indicated by a vector, ft, drawn (in light green) vertically upward along 

the polar axis from the Earth's center, C. The length of 

this vector represents the magnitude of the constant 

angular velocity; the direction of the vector (righthand-thread rule) signifies that the Earth's rotation is 

easterly, as shown by the small circular arrow near the 

upper limit of the polar axis. Because every point on 

Earth rotates in the same direction and at the same 

angular velocity around the polar axis, a similar vector 

is drawn (in light green) vertically upward at each of
32 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

the points P1? P2 , and P3 . Thus, in the same way that 

the vector drawn at point C represents a complete 

statement on the magnitude and direction of rotation of 

the ellipsoidal Earth, so the vectors drawn at Plf P2 , 

and P3 represent complete statements on the same 

magnitude and direction of Earth rotation at these 

points.

It is now a relatively easy step to account properly for 

the Earth's rotation simply by resolving the rotation 

vector into its equivalent components along the x-y-z 

axes of the stable platform. Noting the situation at P1? 

P2 , and P3 (fig. 17), no component will be in the y-axis 

direction as long as the stable platform stays correctly 

oriented. At Pl (on the equator), the entire vector, ft, is 

in the x-axis direction. Thus, for any location on the 

equator, the stable platform needs only to be rotated 

around its x axis at the full magnitude, ft. Similarly, at 

point P3 , as locations are visualized progressively 

nearer the pole, the limit indicates that the entire vector, ft, is in the negative z-axis direction. Thus, the 

stable platform needs only to be rotated around its z 

axis at the full magnitude, ft, in the negative direction.

A general statement can now be drawn, for any inertial navigator location in the northern hemisphere, 

describing how the stable platform must be commanded to preserve its correct orientation despite the 

rotating Earth. The ingredients for this statement 

emerge from the situation at P2 (fig. 17). Note first that 

the Earth's rotation vector, ft, can be resolved into an 

equivalent set of two components, ftX2 and fL , along 

the x axis and z axis (extended), respectively. Note also 

that the angle between ft and the x axis is the latitude, 

4)2. Thus,

ftX2 = ft cos and ft/2 = -ft sin

For any latitude, 3>, therefore, the general statement 

simply requires a constant rate of rotation of the set of 

axes, and, by implication, the stable platform of the 

IMU, so that the component rotational rates are

wx = ft cos

wz = -ft sin 3>,

where wx, wy, and wz signify the component angular 

rates of rotation of the stable platform around the indicated axes.

The above expressions are very similar to those developed from the spherical geometry (for analyzing 

eastward components of change in inertial navigator 

location) if ft is recognized simply as analogous to a 

time rate of change in the longitude, X.

An added comment is appropriate concerning the 

technique for rotating the stable platform of the IMU at 

some stipulated angular rate about any one of its three 

axes; for example, suppose the stable platform is to be 

rotated around its x axis to satisfy the relation

wx = ft cos

From the earlier discussion on gyro and from the schematic (fig. 6) showing the three-gyro array, note that 

the input axis of the x gyro is the axis around which the 

specified rotation must occur. For this to happen, a 

properly proportioned torque must be applied around 

the output axis of the x gyro, which, thereby, causes 

that gyro to rotate (precess) around its input axis. The 

IMU, then, through the appropriate servo loop, discovers that the null balance of the x gyro has been disturbed and, in seeking to reestablish that balance, commands the proper gimbal torque motors to rotate the 

stable platform the desired amount and in the desired 

direction.

For simplicity, the discussion thus far has considered 

only the principal environmental features of an idealized symmetrical and homogeneous Earth. Validity 

has not been jeopardized by this approach, however. 

The steps remaining to achieve a completely rigorous 

analysis of all environmental parameters in the real 

world are definable and can be taken into account in a 

straightforward manner. They have been painstakingly detailed in a number of texts on inertial guidance 

and navigation (see, for example, Britting, 1971; 

Draper, Wrigley, and Hovorka, 1960). Thus, much is 

known about the distribution of gravity disturbances 

and how they influence the accelerometer outputs. 

Also, much is known about the exact configuration of 

the Earth and how it departs from the idealized shape 

of an ellipsoid.

It is hoped that the Earth-environmental stage has 

now been set completely enough so that perspectives on 

the performance dynamics of the inertial navigator can 

be further enlarged. The approach here is to envision 

the system installed in the carrying aircraft, then to 

examine the basic steps needed to accomplish initial 

alinement, and finally to allow the aircraft to be flown 

and to describe how the inertial navigator can develop, 

in a dynamic way, a continous record of the flight path, 

referenced to the onboard set of three coordinate axes 

which, in turn, are keyed to local latitude, longitude, 

and elevation.

To accomplish the initial alinement of the inertial 

navigator, it must be presupposed that the aircraft in 

which it is mounted is at rest over a point on the airport 

parking ramp for which the coordinates of latitude, 

longitude, and elevation can be given. These data may
INSTRUMENT SYSTEM 33

be supplied to whatever accuracy is obtainable from the 

best available local map, as long as the uncertainty in 

location is no more than ±200 ft (61 m) for the horizontal coordinates and ±10 ft (3 m) for the vertical. The 

vertical criterion really comes from the precision with 

which an aircraft altimeter (a special pressure transducer) can be read while navigating enroute to the survey site. As a reference guide, the positional accuracy 

for locating a well-defined point on a representative 

Survey standard 7.5-minute quadrangle map is approximately ±40 ft (12.2 m) for the horizontal coordinates and plus or minus one-half the contour interval 

for the vertical. In nonmountainous terrain, the contour interval is commonly 10 ft (3.0 m), and, thus, the 

vertical coordinate data are accurate to ±5 ft (1.5 m).

The foregoing coordinate data constitute the necessary initial position information for the inertial navigator. The remaining requirements for describing or 

establishing the initial conditions are satisfied by noting that the speed of the aircraft is zero and by manually commanding the x, y, and z axes of the stable 

platform of the IMU to settings within a few degrees of 

the desired tangent-plane orientation. The proper components of Earth rotation, precomputed for the latitude 

given in the initial position data, now are translated 

into precession torques that are applied to the output 

axes of the x and z gyros. The y gyro need not be 

torqued inasmuch as its desired orientation is with its 

input axis perpendicular to a meridian plane of the 

Earth (see fig. 17) which means it then will not sense 

any Earth rotation.

At this point in the procedure, the stable platform 

probably will be out of the intended tangent-plane 

alinement by a few degrees and will be slowly drifting 

more out of alinement by whatever components of 

Earth rotation the misalined gyros sense. Next, the 

three accelerometers are read, and their outputs resolved into components along the x and y axes. The 

amounts by which these components disagree (ideally, 

the disagreements should all be zero when the stable 

platform is level) indicate the tilt remaining in the 

stable platform with respect to the horizontal plane and 

represent the amounts by which it must still be rotated 

around its x and y axes to reach the tangent-plane 

condition. The required amount of rotation, around, for 

example, the y axis is, to a first approximation, the 

arcsine of the x component of the gravity vector, gx, 

divided by the full gravity vector, g. To accomplish the 

leveling, the x and y components of accelerometer outputs are fed through appropriate circuits to the torquers of the y and x gyros, respectively. Precession of 

the gyros, in combination with the action of the gimbal 

servoamplifiers and torque motors, then rotates the 

stable platform to the desired level condition.

Initial leveling may be accomplished in a minute or 

two using relatively high gains in the torquing loops. 

Meanwhile, azimuth misalinement (failure of x axis to 

point true north) will continue to cause the stable platform to rotate around the y axis as the y (east) gyro 

senses a component of Earth rate proportional to the 

misalinement angle multiplied by the latitude. The 

component of gravity along the x axis, measured by the 

accelerometers as the stable platform rotates off level, 

is fed to the z (azimuth) gyro with appropriate gain to 

null the azimuth error. The dynamics of this process, 

called gyrocompassing, are very much slower than for 

leveling. At least 15 or 20 minutes are necessary to 

accomplish the azimuth alinement. Final alinement in 

level and azimuth takes place with the leveling and the 

gyrocompassing loops in operation.

At this juncture, let it be assumed that the inertial 

navigator, as part of the APT system, has successfully 

completed all initial alinement steps. Although stationary in a parked aircraft for the moment, it is ready 

to navigate.

The aircraft engines now are started, and, as the 

aircraft moves along the taxiways toward the departure runway, the accelerometers sense (measure) this 

new movement, as well as any changes in local gravity, 

in terms of the equivalent three component movements 

along the x, y, and z axes of the stable platform. From 

the earlier description of the accelerometers, it should 

be noted that the measurement data actually recorded 

are in units of velocities, v, inasmuch as these instruments have automatically accomplished one integration of the specific force sensed.

The component velocities (ground speeds) vx and vy 

are for the north or south and east or west directions, 

respectively. These are velocities along the Earth's surface or perpendicular to its radius of curvature and to 

the direction of gravity (local vertical) at the alinement 

point.

In a dynamic way, therefore, the inertial navigator 

performance, as it defines changes in the two horizontal coordinates of position, can be described as a closedloop system with two built-in stages of mathematical 

integration. The first stage is built into each accelerometer; the second stage is executed by the on-board 

computer. The mathematical designation for such a 

loop is a "second-order system," and this one is undamped. In the mathematical equation that describes 

the performance of this system, the relative values of 

the coefficients signify that the system is prone to oscillate. For this closed-loop system, if there should be an 

error in the initial conditions as, for example, a slight 

tilt of the stable platform off tangent-plane orientation 

(that is, the z axis not quite coinciding with the plumb 

line of the actual gravity vector as navigation begins), 

then the accelerometers will sense gravity components
34 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

proportional to the tilt magnitude and direction. These 

are integrated once in the sensing process and again in 

the computer; they are approximated by dividing respectively by the northerly and easterly radii of Earth 

curvature and integrating continuously to obtain aircraft position. As successive cycles of events are compared, the change in output of the loop or the loop 

"gain" must be a function of gfR, where R is an appropriate value of Earth radius. This ratio of an acceleration divided by a length, coupled with the idea of an 

oscillating system, leads to the concept of a pendulum 

of arm length, R, subject to gravitational acceleration. 

The period at which such a pendulum would oscillate 

may be shown by simple computation to be 84.4 minutes. In the inertial navigation literature, this is 

termed the "Schuler period," and it characterizes this 

closed-loop system that defines changes in the horizontal coordinates of position. The important point is that, 

although the initial condition errors, as well as some 

other errors that reflect hardware imperfections, cause 

oscillations in the computation of position and velocity, 

the loop exhibits neutral stability; that is, it does not 

"run away" (oscillations steadily increase).

To this point, the discussion has presumed ideal or 

perfect inertial instruments. However, the minor 

asymmetries known to be present in such instruments 

cause errors in alining the IMU. Thus, a bias in an 

accelerometer output causes an error in leveling the 

stable platform or a drift bias in the y (east) gyro causes 

a misorientation in azimuth. To forestall these kinds of 

problems, the final stage in fabricating the IMU is to 

run it through an elaborate one-time program of laboratory calibration exercises. These serve to measure 

and document, very precisely and systematically, the 

magnitude, sign, and, sometimes direction in space of 

such parameters as scale factors, biases, gyro drifts, 

gravity-sensitive unbalances, and mechanical misalinements. Such data are enhanced by additional calibration exercises in the aircraft before each flight, and 

the combined data are stored in the onboard computer 

memory for subsequent operational use as appropriate.

Thus far, the aircraft has simply been taxiing along 

the ground, and the discussion, therefore, has focussed 

first on processing the accelerometer outputs to yield vx 

and vy data, inasmuch as the principal motion is in 

these two horizontal coordinate directions. However, 

the vertical position accuracy specified for the APT system, in its operational surveying mode, is of such a 

high order [±0.5 ft (0.15 m)] that data from a barometric altimeter will not suffice, and the system must be 

capable of determining vertical-coordinate (elevation) 

information as well as horizontal. The mechanization 

of this capability has certain unique characteristics 

and requirements.

After the IMU is initially alined, with the aircraft 

standing in place (zero velocity) over a known point, 

the accelerometers can measure continuously the actual gravity vector. Their respective measurements are 

mutually orthogonal components, integrated once, of 

the specific force vector that acts on the aircraft. These 

measurements are resolved continuously in the computer into components along the x, y, and z coordinate 

axes of the stable platform, but, because, in this case, 

the aircraft is not moving, the resolution yields a 

stream of measured values of integrated specific force 

representing the resultant force along the z axis only. 

These values, which have the dimensions of velocity, 

increase linearly with time because only the gravity 

vector is acting, and they are, thus, the "zero" reference 

for subsequent computations of vertical velocity when 

the aircraft moves. The zero reference or initial value 

for the vertical coordinate of position (the "elevation" 

coordinate) is determined from the known elevation of 

the point over which the aircraft is standing.

When the aircraft moves (taxis), the aforementioned 

accelerometer-measurement and computer-resolution 

processes continue, but now the stream of values that 

represent the changing vertical velocity must be further processed. The appropriate zero reference values 

are subtracted, and the results are integrated over discrete (small) time intervals to yield vertical distances 

that signify changes in the elevation coordinate of aircraft position. As each new elevation coordinate is determined, it must be fed back through a computer program which describes how the gravity magnitude 

changes with altitude. After one integration, each new 

or adjusted gravity value is subtracted from the next 

z-axis component of the integrated specific force as 

derived from the continuing accelerometer measurements5 . The measurement and computation processes 

continuously progress, or cycle, in this closed-loop manner and include compensation for the "Coriolis acceleration" (described later in this section).

Again, the mathematical designation for such a 

closed loop is a second-order system, and it also is undamped. Now, however, an important difference in the 

mathematical equations that describe performance of 

the elevation-coordinate loop, as contrasted with the 

horizontal-coordinate loops, arises from the fact that 

the Earth's gravitational field changes with altitude; 

that is, decreases as altitude increases. The practical 

consequence is that, whereas in the neutrally stable 

horizontal-coordinate loop the errors were oscillatory

5In practice, two models of the gravity field are used, one in real time and the other in 

postprocessing. The real-time model is simply "1 g"; that is, the gravity field of an idealized, 

rotating, ellipsoidal Earth. This model permits the accelerometers to deliver information 

which, to a first approximation, is the desired inertially referenced specific force. The 

postprocessing model is more elaborate, designed to specify departures of the actual gravity 

field from the idealized one.
INSTRUMENT SYSTEM 35

and described with sine and cosine functions, now the 

errors for the elevation-coordinate loop are described in 

terms of hyperbolic sine and cosine functions which 

grow exponentially, doubling in value every 395 seconds. Thus, the mathematical solutions for the performance equations are seen to "diverge," and the 

elevation-coordinate loop is not stable and tends to run 

away. A bias error of only one-millionth g, in determining a new or refined gravity value for a new elevation 

coordinate of position, therefore, can grow in 3 minutes 

to an error of 6 in. in the vertical position.

Because of the foregoing inherent instability problem, the inertial navigator alone is incapable of meeting the stringent positioning requirements. A source of 

position-related measurement data external to the inertial navigator, therefore, must be supplied at regular 

intervals during system operation; this external information is acquired by the "laser tracker," described in 

the section so titled. Enroute from the alinement and 

takeoff location to the field survey site, error growth in 

the elevation-coordinate loop is limited by using data 

from a very precise pressure altimeter to compute and 

reset the vertical coordinate of position. Earth radius 

also is recomputed from these data (accurate to about 

five parts in a million) and used in the horizontalcoordinate loops.

The Coriolis acceleration, mentioned above, deserves 

limited elaboration here. Succinctly stated, this effect 

represents an "apparent" acceleration that arises because the inertial navigator is required to keep track of 

its velocity and position in terms of a three-coordinate 

scheme of reference (latitude, longitude, and elevation) 

that is tied to a rotating Earth. The inertial navigator 

must use accelerometer measurements that are affected by that rotation. Newton's laws of motion applied to the rotating reference scheme define the 

Coriolis acceleration term (see, for example, Britting, 

1971, p. 17-18).

A more meaningful description, couched in terms 

that reflect a real-world phenomenon, hinges upon 

careful definition of where the observer is, and his state 

of motion, when he witnesses the Coriolis effect. First 

imagine the observer to be stationed out in space so 

that he sees the stars as stationary or fixed and he sees 

the Earth, as it appears in figure 17, rotating about its 

polar axis. If an artillery gun at P2 were to fire a shell 

horizontally (tangent to the Earth) in, for example, an 

easterly direction, then the space observer would view 

its initial path of motion as an undeviating straight 

line lengthening eastward. He would mathematically 

describe it, simply and correctly, using a straightforward application of Newton's laws in the presence of 

the Earth's gravity field. Now imagine the observer on 

the Earth at P2 and fire a second shell in the same

manner. The initial path of motion to the earthbound 

observer "appears" to curve steadily upward and to his 

right, departing more and more from the desired easterly direction which, for the rotating Earth, is defined 

by the parallel of latitude through P2 . The observer, 

with recourse only to earthbound reference schemes, 

finds that the complete mathematical description of the 

observed curving path of motion requires an added 

term in the application of Newton's laws. This reconciling term, which accounts for the shell's apparent acceleration upward and to the right of the easterly direction, is the Coriolis acceleration.

Although the foregoing description has, for convenience, featured a projectile in flight, the principle is 

equally apt for aircraft flight. Thus, the dynamic performance of the inertial navigator includes appropriate 

compensation inputs to offset the Coriolis acceleration.

A logical conclusion for this section on performance 

dynamics occurs when the aircraft reaches the departure runway and takes off. The inertial navigator simply continues to perform in the dynamic manner detailed throughout this section. As the aircraft gains 

altitude and normal cruising speed, however, the respective changes in elevation, latitude, and longitude 

rates become far more dramatic than those occurring 

during the taxi operation. Through the onboard computer, these data are processed as a continuous stream 

of position-coordinate information which is not only fed 

into updated computer memory, but also recorded on 

tape to document the aircraft flight path.

External Measurement Requirement

In the very early stages of shaping this instrumentation development project, before the first phase of contract work at Draper Laboratory was formalized, the 

stipulation for achieving position measurement accuracy of ±0.5 ft (0.15 m) in the vertical coordinate direction was recognized as invoking a requirement for updating the inertial navigator at 3-minute intervals. 

The origin of this requirement is found in the mathematical expression that describes how accelerometer 

measurement errors develop in the vertical direction as 

a function of time. It turns out that the time span, t, 

between updates is closely approximated by the familiar relation

s = (V2) at 2,

where s is the 0.5-ft (0.15-m) allowable position error in 

the vertical coordinate, a is the error in measuring 

vertical acceleration, which is relatable to performance 

characteristics of the accelerometers, and t is the time 

interval over which the allowable 0.5-ft (0.15-m) position error develops. A value for a that reasonably typi-
36 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

fies the APT system accelerometer is 1 x 10~6g, where 

g is the standard acceleration due to gravity, or 32.2 

ft/s2 (9.8 m/s2). Substituting in the foregoing approximation and solving for t yields a value of 176 seconds or 

3 minutes.

The update requirement is satisfied at a given instant by feeding into the computational procedures, 

which produce the indicated position data, very precise 

observations of the carrying vehicle's (aircraft's) velocity vector and position in three-dimensional space. 

How to make such observations from an aircraft in 

flight was the real challenge.

At the beginning of the 6-month engineering analysis made by Draper Laboratory, it was believed that 

observations of aircraft velocity could be made at the 

desired 3-minute intervals by the laser profiler if it 

were aimed at and electronically locked onto a retroreflector target at a known ground-control point. By collecting (while locked on) a stream of slant-range, 

azimuth-angle, and dip-angle data, all referenced to 

the orientation (three coordinate axes) and position of 

the onboard IMU stable platform, the laser profiler 

would yield the information needed to compute very 

precisely the aircraft velocity vector.

As the analysis progressed, however, it became apparent that the primary function of the laser profiler 

should not be interrupted to the extent represented by 

the inertial navigator update requirement. Furthermore, a laser-ranging device, which is obliged to operate generally with uncooperative targets (terrain features), should be designed very deliberately and 

exclusively for the purpose. In sharp contrast are the 

design requirements for a similar device scheduled to 

operate only with cooperative targets (retroreflectors). 

And, finally, it was recognized that, for those applications of the APT system where terrain profiling was not 

a requirement, the system should be self-sufficient 

without the laser profiler and be able to perform its 

basic function of carrying a three-coordinate reference 

frame for other ancillary equipment as dictated by the 

particular application. Thus, the idea of a laser ranging 

or tracking device, integral with the IMU, evolved.

Electro-optical, mechanical, and operational details 

of the laser tracker are given in the next section. The 

need for the range and angle data obtainable through 

the operation of such a device relates to the characteristic inherent in any inertial system whereby its indicated position and orientation slowly deviate from the 

true values. The deviation reflects the combined inaccuracies originating in the performance of the three 

gyros and three accelerometers, each one of which 

drifts very slowly.

The practice of aiding or updating an inertial measurement unit is well established and not unique to the 

APT system. An element of uniqueness, however, lies

in the overall accuracy specified for the system, and 

this obviously can be no better than that of the updating technique and equipment.

In essence, the update procedure can be regarded as 

a tripartite affair that involves aiding the inertial 

measurement unit in tracking vertical motion of the 

aircraft with updates from a very sensitive baroaltimeter, while navigating to the survey site; fine calibration 

of the instrument system when it is first flown over the 

survey site; and subsequent updating adjustments to 

the system at no longer than 3-minute time intervals 

throughout the flight mission to hold performance 

within the stipulated precision limits.

The boroaltimetry is accomplished with a pressure 

transducer connected to a suitable static-air-pressure 

port on the aircraft. Transducer details are described in 

the "Laser Tracker" section of this paper.

The calibration exercise requires flying a specifically 

defined path over and around a sequence of three 

retroreflector targets prepositioned on three groundcontrol points, chosen so they are not in a single 

straight-line alinement. The x, y, and z coordinates 

(latitutde, longitude, and elevation) must be determined or known for each control point before the flight 

mission. Although not a requirement, the availability 

of a gravity measurement at each control point would 

simplify the inflight calibration procedure.

Laser Tracker

The manner in which the need for a laser tracker 

component evolved is described in the preceding section. The design of the laser-ranging elements for this 

instrument component benefited from the laboratory 

experiments that shaped the design for the "laser profiler," and evolutionary details on that design are given 

in the section titled "Laser Profiler." The present section, therefore, is limited mainly to the electro-optical, 

mechanical, and operational details of the laser tracker 

as they gelled in the final design. The term "laser" is an 

acronym that signifies "Light Amplification by Stimulated Emission of Radiation."

To ensure that its pointing measurements can be 

accurately related to the stable platform reference 

axes, the tracker two-axis gimbal-support structure is 

mounted in the same housing or case that contains the 

IMU. The arrangement is shown schematically in figure 18, which should be recognized as simply a more 

complete version of figure 10. To avoid gimbal lock for 

all pointing directions downward below the aircraft, 

the tracker outer axis (fig. 18, blue color code) is constructed parallel to the aircraft pitch axis. The tracker 

inner axis (fig. 18, red color code) is always at right 

angles to the outer axis regardless of pointing direction. Arrowheads on the two axes again denote the 

positive-direction conventions.
INSTRUMENT SYSTEM 37

The tracker assembly can be rotated about each axis 

±60° from the position shown (line-of-sight downward). 

The termini of the inner axis are in the gimbal; the 

termini of the outer axis are in the enclosing housing or 

case. The case structure, therefore, provides the fixed 

geometric reference link between the tracker and the 

IMU. This design aids the whole process of searching 

for and acquiring a retroreflector target at a groundcontrol point. Thus, in approaching a target, the 

tracker can be made to scan the ground in raster (sideto-side) fashion. The technique is to point it forward 

along the intended flight path by rotating its supporting gimbal around the outer axis; the desired groundscanning motion is then obtained by simultaneously 

oscillating the tracker assembly itself (lower moment 

of inertia) from left to right around its inner axis.

AZIMUTH AXIS

ELEVATION AXIS

SUPPORT HOUSING

SUPPORT HOUSING 

OR CASE

RESO1VER

AIRCRAFT CABIN FLOOR

FIGURE 18. Support structure for inertial measurement unit and tracker.
38 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

Many of the electromechanical assembly features of 

the IMU are repeated in the tracker assembly. Thus, 

each end of each axis fits into a pair of preloaded ball 

bearings which, in consort with the support-structure 

members, feature an overall structural stiffness capable of maintaining relative positions within microinch 

tolerances. To enhance this capability, the main support housing or case for the whole IMU and tracker 

assembly (fig. 19) is to be made of 6061 aluminum, 

which, for its low weight (density), offers good thermal 

conductivity, high structural and mechanical stability, 

and high resistance to corrosion. The enclosing jacket 

or housing for just the tracker unit itself (fig. 19) is to 

be made of a single piece of heat-treated Ti-6Al-4x 

titanium alloy, which, in addition to the above characteristics, has the added advantage of a thermal coefficient of expansion closely matched to that of the optical 

glass.

Mounted as appropriate in gimbal or case at one end 

of each of the abovementioned two axes is a torque 

motor to drive the tracker around that axis (fig. 18). 

Each of these two motors has a torque capability of 60 

oz-in. (0.42 N-m). Mounted as appropriate in gimbal or 

case at the opposite end (from the torque motor) of each 

of the foregoing two axes is a l:36-speed resolver (fig. 

18). Each measures the angular displacement of the 

gimbal, and the readings are relayed to the onboard 

computer. The net effect of these cited electromechanical features is to assure that the overall tracker pointing accuracy is not degraded from the design accuracy 

of the resolver, namely, about 4 arc-seconds.

A few words are appropriate to interpret the color 

coding used in figure 19. Individual colors are listed 

and explained as follows, proceeding generally from 

outermost to innermost physical features:

Blue Main support housing or case for 

entire IMU and tracker assembly,

Pink Elevation gimbal structure, 

Green Stable platform structure and 

internal electronics packages, 

Red Gyros, 

Silver Accelerometers, 

J Gold Gimbal structure, and 

[Lavender Part of tracker unit that moves 

on inner axis; also inset, at 

lower right, which shows 

beam splitter, range receiver, 

and quadrant photodiode assemblies.

The sweep angle of the tracker (±60° from the vertical) had to be obtained without letting any part of the 

assembly protrude beyond the bottom exterior surface

IMU

Tracker

of the aircraft cabin. This specification is met by constructing a floor well of appropriate dimensions, 

mounting the tracker as deep therein as possible, and 

adopting a coaxial and folded optical lens system. Layout of the tracker optics is shown in figure 20. From 

this layout, the laser-beam (green band, fig. 20) path is 

traced readily from source through transmitter lens 

assembly, to deflecting mirror, and to deflecting prism 

centered and cemented on the back side of the objective 

lens. At this point, the laser beam is outbound toward 

a retroreflector target secured on a ground-control 

point. After reflection, the laser beam returns inbound 

(light-blue band, fig. 20) through the same objective 

lens, to the deflecting alinement mirror, and through 

the beam-splitter assembly (medium-blue band, fig. 20) 

for delivery to a quadrant-photodiode-receiver 

assembly and a range-receiver assembly; the former 

assembly receives 90 percent of the return signal, and 

the latter, 10 percent. The 4.5-in. (114-mm) diameter 

objective lens [11.902-in. (302.3-mm) focal length] was 

the largest that could be used because of the 24-in. 

(0.61-m) diameter clear limit originally placed on the 

aircraft cabin-floor well.

To inhibit moisture condensation in the tracker optics, two silica-gel capsules occupy a cavity designed 

into the tracker unit housing. The capsules are packaged in a lint-free nondusting material, and, by choice, 

the gel does not feature any dye to indicate moisture 

content, thereby forestalling any opportunity for adversely affecting a lens coating.

The variable density optical filter, shown in the path 

of the outgoing laser beam, is operated by an automatic 

gain control loop which monitors the intensity of the 

return laser beam as the sum of the four quadrant 

pulse magnitudes. Thus, as the return signal weakens 

or strengthens, owing, for example, to changes in attenuation as a function of range to the retroreflector, the 

servo-drive motor rotates the continuously variable filter just enough to strengthen or weaken the outgoing 

beam the needed amount. The intent is to maintain a 

constant sum for the four quadrant return-signal 

pulses.

The range receiver assembly features a silicon 

avalanche photodiode detector and transimpedance 

preamplifier. It senses the return-signal pulse from 

which range to the retroreflector target is ultimately 

determined.

The laser beam is generated by a pulsed gallium 

arsenide (GaAs) transmitter that is similar to the one 

designed for the laser profiler. Pulses are produced at a

FIGURE 19. The inertial measurement unit, laser tracker, 

gimbal-support structure, and housing.
INSTRUMENT SYSTEM 39
40 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

rate of 3.2 kHz in the near infrared at a wavelength of 

904 nanometers (nm), with 25 watts (W) peak power, 

subnanosecond rise times, and widths of 40 nanoseconds. These large widths (four times greater than the 

profiler pulses) assure that more return-signal energy 

will be fed to the quadrant photodiode assembly 

(fig. 20) and to the tracking electronics. The laser beam 

meets Federal safety standards in the following manner:

1. In the laboratory. Eye safe for all distances greater 

than 18 in. (0.46 m) when using a suitable attenuation filter and limiting exposure time to a 2-hour 

maximum.

2. In flight. Eye safe for all distances greater than 

892 ft (304 m). This type of laser beam permits 

range determinations well within the stipulated

precision of ±0.5 ft (0.15 m) over distances as 

great as 6,000 ft (1.8 km), which is the maximum 

planned operating slant range to a retroreflector 

target. Divergence of the laser beam is 6 to 7 milliradians (mrad) and the fields of view of the 

quadrant-photodiode and range-receiver assemblies, respectively, are 8.3 and 2.6 mrad. 

The quadrant photodiode receiver assembly senses 

the return signal as four independent pulse magnitudes related to the sectors above, below, left, and right 

of the quadrant dividers. If the tracker boresight 

catches a retroreflector target off center, the target image will be unevenly distributed in the four quadrants, 

and the four pulse magnitudes proportional to the 

sighting error in each quadrant, and the corresponding 

control currents drive the two torque motors (see fig.

ALIGNMENT 

MIRROR

LASER

TRANSMITTER 

ASSEMBLY--

LENS

MIRROR

DC TORQUE MOTOR

VARIABLE DENSITY 

FILTER DISC

RANGE 

RECEIVER 

ASSEMBLY

BEAM SPLITTER

QUADRANT

PHOTODIODE

ASSEMBLY

FIGURE 20. Optical layout for laser tracker.
INSTRUMENT SYSTEM 41

18) to keep the target image centered in the tracker 

field of view.

The discussion given in the section, "Laser Profiler," 

refers to a key mathematical expression termed the 

"range equation" which is equally important in this 

section in arriving at suitable design choices for the 

laser transceiver. Although the following discussion of 

the range equation closely parallels that given in the 

"Laser Profiler" section, there are significant differences that relate primarily to the laser beam ranging 

on a specially designed cooperative retroreflector 

target instead of the uncooperative land-surface target. 

Thus, for the laser tracker, the range equation has the 

form:

Pp = P<

T1T2pA1A3e-^R 

1 A2A4

(3)

where PR = power, receiver, 

Ps = power, source,

T! = transmission coefficient, source optics, 

T2 = transmission coefficient, receiver optics,

p = reflection coefficient, retroreflector, 

A! = area, retroreflector, 

A2 = area, light beam "footprint" at retroreflector,

A3 = area, receiver lens,

A4 = area, light beam return "footprint" at receiver, 

a = attenuation coefficient, atmospheric path,

and

R = range.

The interrelations among the foregoing parameters 

are ilustrated schematically in the three parts of figure 

21. These parts, studied in sequence from left to right, 

show (A) the significant geometric features of the laser 

light path from transmitter (source) to retroreflector 

target and back to receiver, (B) the critical points in the 

flow of light energy from the lasing source to the target 

and return, and (C) the mathematical description of the 

principal losses in energy flow which allows a rigorous 

statement to be written (eq 3) that shows the net 

change in energy level between laser source and 

receiver. Obviously figure 2 LA is schematic in the 

sense that, in the real tracker instrument, the field of 

view is coextensive (not side by side) with the field 

irradiated.

To arrive at the particular form of the range equation 

given in figure 21, note from the geometry shown in 

part (A) that the radius of the circular footprint of the 

laser beam at the retroreflector is a function of the 

constant beam divergence angle, a, and the range, R. 

Thus, the footprint area, A2, is determined from the 

relation,

A2 = K2TrR2,

where K2 is simply the appropriate trigonometric function of a used with R to compute the footprint radius. To 

develop a similar expression for A4, the area of the 

circular footprint of the returning laser beam at the 

receiver, consider first that the retroreflector is designed to have a high reflection coefficient and a 

shaped surface such that virtually all the laser beam 

that is intercepted is reflected straight back along the 

line-of-sight path to the source. In effect, therefore, the 

retroreflector becomes a new point source for the laser 

beam, with its own constant divergence angle, a'. 

(Ideally, a' should be nearly equal to, but no less than, 

a.) Thus, the radius of the circular footprint at the 

receiver is a function of a' and R, and the area, A4, is 

determined from the relation,

A4 = K4TrR2 ,

where K4 is the appropriate trigonometric function of 

a'. By substituting the foregoing equivalences for A2 

and A4 in equation 3, the range equation for the laser 

tracker has the desired form,

Pp = P<

T1T2pA1A3e-2^

K5TT2R4

(4)

where K5 is simply a new constant that stems from the 

product of K2 and K4.

The tracker may be operated in either a search mode 

or a track mode. In the search mode, when the tracker 

must search until it acquires a retroreflector target, the 

tracker gimbals are positioned automatically by computer program in much the same manner as described 

in the last paragraph in the section titled "Electronics." 

Now, however, the appropriately processed gimbal 

angle data are sent to the tracker gimbal servoamplifiers to drive the tracker torque motors. Included in the 

computer program for this pointing mode is continuing 

iterative calculation of the commanded tracker gimbal 

angles to maintain the line of sight toward the expected 

position of the retroreflector target. This takes into account the known target location, as well as the present 

aircraft position, attitude, and velocity, and then superimposes a tracker spiral scan motion that slowly 

opens up, in a step-by-step programmed way, around 

the initial, most probable target direction. At the maximum intended flight operating altitude of 3,000 ft 

(920 m), the spiral scan will sweep the tracker footprint 

over a terrain area that totals about 4.3 x 104 ft2 

(4.0 x 103 m2) in 20 seconds. Normally, this is more 

than sufficient to find the retroreflector target. However, if the target is not found and is determined to be
LIGHT PATH IN SPACE 

(GEOMETRY)

TRANSMITTER 

LENS

LASER BEAM

DIVERGENCE

ANGLE, a

LASER BEAM 

FOOTPRINT, A2 A^

(A2 =

RANGE, R

REFLECTED BEAM

DIVERGENCE

ANGLE, a'

(RETROREFLECTOR) ' A, 

) (A4 = K47tR2)

e

FLOW OF LIGHT ENERGY 

(FIRST LAW, THERMODYNAMICS)

(RETROREFLECTOR)

OUTGOING PULSE

MONITOR DISPLAY 

RETURN PULSE

MA

D

RAN

PR = P

(SEE EQUA

TIME

FIGURE 21. Analysis of the range equation for the laser tracker.
INSTRUMENT SYSTEM 43

still within range, then the computer program commands the tracker to enter a side-to-side oscillatory or 

raster scan. If this broader search mode were to persist 

the full program length to 40 seconds, then the tracker 

footprint would sweep over a terrain area that totals 

about 7.8 x 105 ft2 (7.2 x 104 m2).

When the tracker receives a return signal from the 

retroreflector, indicated when the total level of returned energy exceeds a preselected limit, an "acquired" flag is set by the signal processing electronics, 

and the tracker then is switched to the track mode. 

Now the outputs of the quadrant photodiode are used to 

generate signals that are sent through the appropriate 

tracker servoamplifiers to the respective torque motor 

(fig. 18) to keep the tracker line of sight centered on the 

retroreflector target. During this tracking-mode operation, readings of the IMU and tracker gimbal angles, as 

well as the target range, are sent to the computer at a 

200-Hz rate. This stream of data provides a complete 

time-history of the aircraft motion vector with respect 

to the retroreflector while it is in view (typically 30- 

60 seconds). In addition, simple curve fitting programs 

allow the computer to extract velocity and acceleration 

information from the data. Should tracking contact be 

lost, the spiral scan submode is reinvoked automatically to reacquire the target within 5 seconds.

The tracker and the profiler operate in conjunction 

with a laser signal processor and time-interval counter. 

The former contains the master quartz-crystal oscillator (clock) for the whole APT system. Its frequency is 

19.6608 megahertz (MHz), and all APT subsystem frequencies are integral submultiples thereof.

A diagram to illustrate the interrelations among the 

principal frequencies is given in figure 22. The master 

crystal is electric oven controlled to a long-term frequency stability of better than 1 in 108 parts. This kind 

of stability is more than academic because the master 

oscillator paces the time-interval counter which ultimately must achieve resolution of better than a 

nanosecond (1 x 10~9 second) if laser ranges accurate 

to one-half a foot (0.15 m) are to be extracted from the 

processed measurement data.

The laser-signal processor generates trigger pulses 

(under control of the time-interval counter) for the 

tracker and the profiler laser transmitters and delivers 

time-multiplexed duration pulses, which become the 

yardstick for determining target range, to the timeinterval counter. The time-interval determinations are 

sent to the onboard computer for conversion to actual 

target ranges.

An electronic refinement in the art of measuring the 

round-trip traveltime of the laser pulse greatly sharpens the ultimate conversion into target range. This refinement, termed a "constant-fraction discriminator," 

is designed to cope with the fact that the returning

laser pulses will range greatly in size. If the timeinterval counter were simply set to make its measurement when the return pulse reached an arbitrary and 

preset voltage level, then this would ignore the known 

variations in pulse size and introduce unwanted errors 

in the range determinations. The way around this 

dilemma is to note that the known time constant of an 

electronic circuit tends to remain stable, and, thus, the 

instant at which the laser return pulse will reach twothirds of its peak value is predictable. The constant 

fraction discriminator, therefore, is built to trigger the 

time-interval count always at the same fraction of 

pulse amplitude, and, thus, all range determinations 

are made from the same relative position on the curve 

that defines the pulse shape. In the tracker, these determinations are made on the leading-edge side of the 

pulse.

The time-interval-counter unit chosen for the APT 

system is the Digitec Model 8330. It is packaged in a 

metal box 5V4 in. (0.13 m) high, 17 in. (0.43 m) wide, 

and 163/4 in. (0.42 m) deep and can be mounted in 

standard light-alloy racks. The unit weighs 21.5 Ib 

(9.8 kg), and its power requirement is 100 W, delivered 

as single-phase current at 113 volts (V) and 400 Hz. 

When it was purchased (in the late 1970's), this off-theshelf commercial unit was unique in its capability for 

achieving a one-shot resolution of 0.1 ns and an accuracy of approximately 0.3 to 0.4 ns. Through timeinterval averaging, however, the resolution can be 

increased to 1 picosecond [ps (0.001 ns)]. As used in 

this part of the APT system, the unit averages 128 

measurements. Thus, its operating resolution will be 

0.4 ns/Vl28, or 0.03 ns.

The time-interval counter unit has its own internal 

voltage-controlled oscillator which operates at 10 MHz. 

Inasmuch as this is synchronized with the master 

quartz-crystal oscillator at 9.8304 MHz (see fig. 22), the 

proportionality constant 10/9.8304 must be entered appropriately into the final range-measurement computations. These computations also must include adjustments for the average refractive index of air over the 

laser light path, based on inflight outside air temperature and static pressure measurements at the aircraft.

To collect the need measurements, a temperature 

transducer and a static-pressure port for a pressure 

transducer are installed at suitable outside points on 

the aircraft. The pressure transducer chosen is a Rosemount Model 1201F1, scaled to provide a DC electrical 

signal in the range of 0 to 10 V, over a range in atmospheric pressure from 17.72 to 32.48 in. (450-825 mm) of 

mercury at 0°C (60,000-110,000 Pa). A 12-bit analogto-digital (A/D) converter transforms (digitizes) the DC 

signal over the 10-V range. Thus, the least significant 

bit corresponds to a pressure change of 0.007 in. 

(0.18 mm) of mercury (24.4 Pa).
44 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

MASTER

CLOCK

QUARTZ

CRYSTAL

OSCILLATOR

19.6608 MHz 

(PERIOD = 50.86 ns)

1.2288 MHz 

(813.80 ns)

614.4 kHz^ 

(1.63^8) ^

307.2 kHz 

(3.62 ns) ^

(6.51 us) "* 

76.8 kHz ^

(13.0 nS)

38.4 kHz ^ __ 

(26.0 ns) ^

19.2kHz -^

(52.1 ns) ^ 

9.6 kHz __

4.8 kHz 

(208^8) ^

2.4 kHz 

(417nS)^ 

1.2 

(833

200 

(5

\

t

©

kHz ^/

\ '

-6

Hz __ fc

T!S) * 

1 '

: 2* -

f

. a

(4.88 ns)

EXPLANATION

KEY TO REFERENCE FREQUENCIES:

1. GYRO WHEEL SUPPLY 

2. GYRO AND ACCELEROMETER 

MAGNETIC SUSPENSION 

3. GYRO AND ACCELEROMETER 

SIGNAL GENERATORS 

4. LASER TRIGGERS 

5. ACCELEROMETER WHEEL 

SUPPLY 

6. GIMBAL RESOLVER 

7. GYRO TORQUE MOTOR 

SUPPLY 

8. IMU MULTIPLEXER AND 

TIME LABELS FOR TV 

FRAMES

/ MASTER 

J TIMING 

\ SIGNAL TO 

\COMPUTER

12.5 Hz 

(80 ms)

nnon i

t

Q) ©

©

© ©

__ ^ 102.4 kHz 

(9.76 ns)

^ 51.2 kHz

^ (19.5 ns)

__ ^ 25.6 kHz 

(39.1 ns)

12.8 kHz

__ ^ 6.4 kHz 

(156 ns)

^ 3.2 kHz

__ ^ 1.6 kHz 

^(625 ^3)

^ 800 Hz

__ k. 400 Hz 

(2.50 ms)

^ 200 Hz

^ (5 ms)

__ ^.100 Hz 

^'(lO ms)

^ 50.0 Hz

^ (20 ms)

__ _ 25 Hz 

^"(40 ms)

12.5 Hz 

(80 ms)

FIGURE 22. Principal reference frequencies used

The temperature tranducer chosen is a Rosemount 

Model 101F probe and Model 510B signal-conditioning 

amplifier, scaled to provide a DC electrical signal in the 

range of 0 to 5 V over a range in air temperature from 

-50°F to +150°F (-46°-66°C). A 12-bit A/D converter 

digitizes this signal also, and the least significant bit

in the aerial profiling of terrain system.

corresponds to a temperature change of 0.195°F 

(0.108°C).

The capability for the tracker to be rotated ±60° from 

the vertical, around each of its two orthogonal axes, 

means that its field of view is the land-surface area 

embraced by the base of a right circular cone whose
INSTRUMENT SYSTEM 45

apex angle is 120°, whose apex is at the intersection of 

the two tracker axes, and whose height is the aircraft 

altitude above the land surface. For the tracker to capture, lock onto, and track a retroreflector target positioned on a ground-control point, imagine the preceding cone to be inverted, and its apex placed instead at 

the control point. As long as the aircraft carrying the 

tracker remains within the volume of airspace bounded 

by this inverted cone, it is physically possible to see the 

target, provided, as already stated, the retroreflector 

has the capability to reflect the intercepted part of the 

laser beam straight back along the line-of-sight path to 

the source. The geometry of the situation is shown in 

figure 23, which has been drawn for the maximum intended aircraft operating altitude of 3,000 ft (915 m). 

For this flight altitude, therefore, the general approximation holds that the aircraft must be sensibly within 

a 1-mi (1.6-km) radius of the retroreflector target for 

any successful tracking activity.

To be visible from all the flight directions diagrammed in figure 23, an ideal retroreflector target 

should be a hemisphere about 1 ft (0.3 m) in diameter. 

To give it the proper retroreflecting capability, the best 

promise for combining accuracy and low cost appears to 

lie in mounting individual molded-plastic retroreflectors, each one of which is about 3/4 in. (19 mm) in diameter, all over its surface.

Until the abovementioned ideas can be thoroughly 

researched and the prospects for commercial fabrication in quantity lots evaluated, the Survey will use its 

own inhouse designed and fabricated retroreflectors. 

The basic elements are four commercial solid-glass,

corner-cube retroreflectors, each having a minimum 

clear aperture of 3V2 in. (8.9 mm) and a maximum 

angular deviation, between the incoming and returning light beams, of ±30 arc-seconds. The four basic 

elements are cemented and clamped to an aluminum 

circular base plate in the attitudes shown in figure 24. 

The in-between spaces are filled, to some convenient 

convex contour, with black silicone-rubber caulking, so 

that, when stationed in the field, little or no opportunity exists for any form of precipitation to collect. The 

base plate is 7 in. (18 mm) in diameter and is provided, 

on a smaller diameter circular plate secured on its underside, with a threaded center hole so that it can be 

mounted on a standard surveyor's tripod. A bullseye 

bubble atop the base plate provides for leveling.

The laser tracker discussion is not complete without 

a brief description of the physical properties of the glass 

window (fabricated more like a large "lens") installed 

in the aircraft cabin floor, through which the laser 

beam must pass. The principal constraints used to identify the most suitable glass are as follows:

1. Stipulated maximum rectangular clear opening, 

27.0 in. (0.686 m) x 33.25 in. (0.845 ml,

2. Maximum air-pressure differential that can be expected to occur between aircraft cabin and the 

outside world, less than 0.5 lb/in.2 (350 kg/m2),

3. Maximum air-temperature differential that can be 

expected to occur between aircraft cabin and the 

outside world, about 75°F (42°C),

4. Maximum window deflection that can be tolerated 

due to any type of loading, less than 0.003 in. (0.08 

mm),

I ,

3,000 ft 

(920 m)

(MAXIMUM FLIGHT 

ALTITUDE)

LAND SURFACE RETROREFLECTOR 

FIGURE 23. Cone-shaped air space within which the tracker can see a retroreflector target

LAND SURFACE
46 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

5. Maximum resistance to corrosion from chemicals in 

the air,

6. Highest transmission coefficient for near infrared 

light,

7. Minimum light-beam distortions due to any residual internal stress, and

8. Minimum number of gas-bubble or striae inclusions

per kilogram of mass.

These constraints led to the choice of Schott B-270 

white opthalmic crown glass to be drawn (as opposed to 

cast) and fire polished in several sheet blanks 28 in. 

(0.711 m) x 34.5 in. (0.876 m) x 0.66 in. (0.017 m). The 

weight of a single blank is about 60 Ib (27 kg). Pertinent physical properties of this glass are as follows: 

Mechanical/thermal

Density is about 1.5 oz/in.3 (2,550 kg/m3) 

Modulus of elasticity is 10.4 x 106 lb/in.2 [72 gigapascals (GPa)]. 

Mean coefficient of expansion6, for the range 20°C to

300°C is (93.3 x 10-7)/°C. 

Optical 

Coefficient of transmission is 92 percent (antireflection

coating will increase this to 98- to 99-percent region). 

Index of refraction for near infrared wavelengths is

1.5298.

Specifications drawn for finishing the glass blank 

that is installed in the aircraft typify the fact that the 

so-called window must really function as a lens. The 

finishing operations included the following:

1. Polishing to stipulated uniform thickness (tolerance 

±0.5 mm and wedge angle less than 4 arcseconds) and surface flatness (both sides),

2. Hounding the four corners of the rectangular sheet 

to a 1-in. (25.4-mm) radius of cuvature,

3. Beveling all edges of the sheet, and

4. Antireflection coating of the sheet (both sides) for 

904-nm wavelength unpolarized light.

Computer

In the array of equipment components that constitutes the APT system, the electrical device that ties 

everything together into a smoothly functioning dynamic whole is the onboard data-processing digital 

computer. This has been partly recognized in the general definition already given for the inertial navigator. 

The objective in this section, however, is to detail more 

completely the manner in which the digital computer 

serves as the central communicator and controller for 

the flow of data, instructions, and control signals 

among all APT system components.

6 A dimensionless ratio of change in length to original length, for a specified temperature 

change.

"S

TJ 

>.

3 

1/3O

C/5

<u 

'o

TJO

CL, O
INSTRUMENT SYSTEM 47

Physical Attributes

The following paragraphs give a description of pertinent physical characteristics and capabilities of the 

particular digital computer selected for the APT system, namely, the Norden PDP-11/70M. The "M" in the 

nomenclature signifies that it is the fully militarized, 

or ruggedized, model suitable for flight application. Its 

structure is diagrammed schematically in figure 25.

The bulk of the Norden computer is packaged in four 

metal boxes, whose dimensions, total weight (mass), 

and electric power requirements are given in table 1. 

The computer control panel is housed separately in a 

fifth box (see table 1). Figure 26 is a photograph of the 

several boxes which, for the desired flight applications, 

can be shock mounted easily in standard light-alloy 

racks secured inside the aircraft cabin. For its ruggedness and compact size, the Norden flight computer is

CO 

LLJ

o

uj 

o

APT SUBSYSTEMS

EXPLANATION

 NUMBERED INPUT/OUTPUT PORT OR CHANNEL 

**CIRCUIT BOARDS

DMA, Direct Memory Access Interface 

GPD, General Purpose Digital Interface 

AS, Asynchronous Serial Interface 

LPC, Line Printer Controller

FIGURE 25. Functional structure of the aerial profiling of terrain system computer with connections to subsystems 

and peripheral devices.
TABLE 1. Package size and power requirements for the aerial profiling of terrain system computer and peripherals

ITEM

I/ 

Norden computer

Computer control panel

Plasma display panel

Line printer

Magnetic tape controller

£/ 

Magnetic tape recorder

TOTALS

Height

in. (m)

7-3/4 (0.20) 

10-1/2 (0.27)

14 (0.36)

9-1/2 (0.24)

10-1/2 (0.27)

7 (0.18)

Width

in. (m)

10-1/8 (0.26) 

19 (0.48)

13 (0.33)

16 (0.41)

17 (0.43)

15-1/4 (0.39)

Depth

in. (m)

19-1/2 (0.50) 

4 (0.10)

12 (0.30)

19 (0.48)

18 (0.46)

14-1/2 (0.37)

Mass

Ib (kg)

b/ 

236 (107)

6 (3)

54 (24)

77 (35)

63 (29)

52.5 (24)

d/ 

541 (246)

Electric 

power

watt

M500

275

750

200

200

d/ 

3125

Electric 

Single p

Single p

Single p

Single p

Single p

a/ Four boxes.

b/ Total values, four boxes.

£/ Two identical units.

d/ Includes two tape recorders,
INSTRUMENT SYSTEM 49

FIGURE 26. Packaged flight computer for the aerial profiling of terrain system.

extremely versatile; only a few of its more obvious features can be described in the following paragraphs.

The structure of the Norden computer, as diagrammed schematically in figure 25, prominently features a single, high-speed, multichannel bus, termed 

the UNIBUS, which allows most of the computer components, peripheral devices, and APT subsystems to 

communicate with each other over its 56 individual 

(parallel) lines. The UNIBUS can be likened to a telephone party line because of its eight numbered input/ 

output (I/O) port connections or channels and its several direct connections to other components.

Paramount among the direct-connected components 

is the "central processor," which occupies a key control 

position in the computer structure. It acts, in part, like 

a central telephone exchange and controls the allocation of time for the communication traffic on the 

UNIBUS. It also performs arithmetic and logic operations as demanded in the decoding and execution of 

instructions. These functions relate to manipulating 

and controlling the APT system through all its intended operational modes.

The central processor is built to recognize and operate at any one of eight distinct priority levels. The 

highest priority is given, for example, to accelerometer 

measurements (see IMU in fig. 25) needed in the realtime computations used to navigate the aircraft. The 

lowest priority is exemplified by a command to print 

out data on the line printer.

The central processor is a key component in other 

important ways that include execution of programs 

that have been stored in memory. Execution can entail 

arithmetic operations, logic operations, or manipulations of data. Arithmetic operations are self evident; 

logic operations involve carrying out instructions; and 

data manipulations involve, for example, moving specified data from one place to another in the computer.

Performance of the central processor is expedited by 

an internal set of 16 high-speed, general-purpose registers in which words (data or instructions) are stored 

temporarily. These help to speed up the computer operation by reducing the traffic of data moving into and 

out of the computer memory.

Operating speed of a computer is characterized by its 

so-called throughput rate, or the number of bits per 

second that it can process. In the Norden, this rate is 

3 x 106 bit/s, and, thus, a basic arithmetic computation, such as the addition of two numbers, can occur in 

about 350 ns.

Allusions already have been made to the computer 

memory (see fig. 25) which provides a more permanent 

form of storage for data and instructions. Memory capacity is an important property, and the Norden computer was chosen partly because it could store 128K 

words, where K signifies 210, or 1024. This assures 

ample reserve capacity beyond the normal loading the 

APT system will put on the computer memory in the 

contemplated flight missions. Furthermore, this
50 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

ruggedized computer model features a nonvolatile 

memory which means that the stored information is 

not lost under electrical power failure.

Positioned between the central processor and the 

main memory (fig. 25) is a very high-speed memory 

unit, termed a "cache," capable of holding 2,048 bytes 

at any given point in time. It augments the main memory in a very dynamic way. Whenever the processor is 

instructed to store words in memory these words are 

written to both the cache and the main memory. The 

premise is that the next step in the instructions may 

call for the processor to use some of these same words 

which occurs a very high percentage of the time. Storage in the cache memory is in constant flux, and the 

newest incoming words simply overwrite (replace) the 

previous words.

The central processor is aided in its arithmetic operations by the "floating-point processor" (see fig. 25), 

which contains six 64-bit registers. These essential elements enable calculations ranging from very large 

numbers (up to about 1038) to very small numbers 

(down to about 10~39). Without this capability, a register, with only a fixed number of bit positions available, 

would be limited to processing only those decimal numbers that fell within a relatively narrow size range of 

large or small.

Successful performance of the APT system, as described in previous sections, relies not only on extremely precise measurements by the gyros and accelerometers but also on real-time and very precise 

computations based on those measurements. The measured quantities are quite small and may involve, for 

example, trigonometric functions whose arguments approach zero for tiny angular displacements near 0° or 

90°. The floating-point processor performs with sevendecimal-place accuracy in single-word calculations. It 

also can perform double-word calculations (so-called 

"double precision" arithmetic operations with doublelegnth numbers) and, in so doing, achieve 17 decimalplace accuracy. Normally, floating-point-processor capability can be provided through either the hardware 

or the software. The real-time constraints imposed by 

the APT system, however, preclude the software approach. The floating-point processor runs independently of and concurrently with the central processor. 

This time-saving computer feature is an invaluable aid 

in coping with the overall workload.

Next among the components that are directconnected to the UNIBUS (proceeding from left to right 

in fig. 25) is the computer control panel (see also 

fig. 26). This unique feature provides a direct link between operator and computer. Thus, the operator may 

preset the knobs and toggle switches, shown in figure 26, to run the computer in a completely automatic, 

semiautomatic, or completely manual mode. The last

two modes commonly are used to test and diagnose 

computer operation; normal operation is fully automatic. The several rows of indicator lights display (in 

binary form) the contents of selected registers and 

memory and, thereby, allow the operator to observe, in 

part, the internal status of the computer. Furthermore, 

the operator can enter his own choices of data or instructions directly into selected registers simply by appropriately setting switches and knobs.

One particular indicator light on the control panel is 

worthy of special note. This light comes on when a 

parity check error is found. The Norden computer uses 

parity checks extensively to ensure that all data and 

instructions are moved to and from the correct addresses in memory. Such extensive checking is essential because erroneous transmission of even a single bit 

could change a tiny number into a huge one (10~9-109, 

for example) or cause the instruction execution sequence to jump to some unintended address, with 

catastrophic consequences for the proper functioning of 

the computer and the APT system. The parity check 

involves computing, transmitting, and storing an extra 

check bit, the parity bit, along with each word of data. 

The parity bit is set to 0 or 1 in such a way that the total 

number of 1 bits in each word is always even. If, in 

subsequent processing actions, an odd number of 1 bits 

occurs, it labels the particular word as having been 

transmitted incorrectly. Special registers in the central 

processor hold information on the occurrence of parity 

errors so that corrective action can be taken automatically. The usual response is simply to reexecute the 

failed instruction.

Subsystems And Data Interfaces

As the review of components direct-connected to the 

UNIBUS continues (still proceeding from left to right 

in fig. 25) the eight numbered I/O ports or channels will 

be described, and the first four link the computer with 

the four principal subsystems into which the whole 

APT system may be divided. These subsystems must 

bear the brunt of whatever earth science measurements the APT system is commissioned to collect, and 

success will ride upon virtually continuous, high-speed, 

two-way communication between them and the computer. Thus, channel numbers 1 through 4 provide the 

data interfaces between the fundamental measuring 

instruments and the computer.

Of the four remaining channels, one establishes the 

means for the APT system operator to communicate 

with the computer, one provides a way to convert selected computer output into hard copy by means of a 

line printer peripheral device, and two will be held as 

spares and not used for the present.

Taken in numerical sequence, channel number 1 

(fig. 25) provides direct, high-speed access to the com-
INSTRUMENT SYSTEM 51

puter memory without interrupting any computations 

that may be in progress. The computer circuit board 

through which electrical connection is made is called a 

Direct Memory Access (DMA) interface. In essence, 

any APT subsystem data or information which should 

be processed in virtually real time are funneled by the 

electronics box (fig. 25) into channel number 1. Examples of such input data are gimbal angle readings from 

the IMU and tracker, velocity readings from the accelerometers, range measurements from the tracker 

and profiler, and mode status of the entire system. 

These input data are processed promptly in the computer, and the needed output data sent back to the 

appropriate subsystem. Examples of such output data 

are torque commands for the x, y, and z gyros in the 

IMU.

Channel number 2 (fig. 25) provides a General Purpose Digital (GPD) interface. A number of analog signals, emanating from the several APT subsystems, 

must be converted to digital format in the electronics 

box (fig. 25) before being multiplexed (transmitted according to a timed sequence) into the computer. Examples of such input analog data are voltage and temperature sensors inside the IMU (generating 

"state-of-health" signals) and certain sensors for the 

operating environment, such as outside air temperature and static pressure. Normally, these data simply 

are converted from analog to digital format, fed 

through the computer, and tape recorded at specified 

intervals. However, the electronics box has the capability to decode directly certain results of the A/D conversions and to generate an alarm if system safety is 

threatened. If future needs arise, this A/D channel can 

accommodate additional subsystem devices.

Channel number 3 (fig. 25) provides a second GPD 

interface. A number of digital outputs (signals) from 

the computer need to be routed to appropriate subsystems and at those destinations converted into analog 

format. Such traffic flows through this channel, and 

decoding for the particular destination is done in the 

electronics box (fig. 25). Examples of the traffic are the 

signals being directed to the servoamplifiers for the 

IMU and tracker gimbal torque motors and to the aircraft autopilot for steering. Digital data passed to the 

computer through this channel include the time-of-day 

code from the master clock (see "Laser Tracker" section 

for clock description). Should the need arise, this channel can accommodate even greater signal traffic if 

other devices that require D/A conversions are added.

Channel number 4 (fig. 25) provides the third GPD 

interface. This channel is the status and command communication link between the computer and the logic 

circuitry in the electronics box (fig. 25) that concerns 

the various modes in which the APT system is to operate. In effect, the electronics box contains two registers,

one termed "Status Register" and the other "Command 

Register." At all times, the Status Register contains a 

16-bit word that indicates the particular mode in which 

the APT system is operating. The computer can be instructed to query this Register and to read the current 

16-bit word, thereby ascertaining the current APT system mode status. If a change in mode status is to be 

effected, the computer will transmit,, through this 

channel, the appropriate 16-bit word to the Command 

Register for implementation.

Channel number 5 (fig. 25) provides a fourth GPD 

interface that is held as a spare in the APT system.

Channel number 6 (fig. 25) affords a two-way link 

between the APT system operator and the computer 

through what is called an Asynchronous Serial (AS) 

interface. The term signifies simply that digital data in 

serial format (a sequence of bits) can be exchanged 

between the computer and the keyboard-plasmadisplay unit in a manner that is not preprogrammed for 

a regular and fixed time interval. The exchange can be 

initiated by the operator or the computer and will occur 

whenever the priority of the particular message can be 

fitted in behind existing higher priority computer traffic. The computer circuit board, through which the electrical connection is made, is labeled AS, and it is bidirectional inasmuch as it feeds into the computer the 

signals from the operator's keyboard and then returns 

the computer's output responses which drive the operator's plasma display. Through this channel, therefore, 

the operator can query the computer and then view the 

responses on the plasma display, as to the status (state 

of health or readiness) of selected APT subsystems or 

the degree of progress through selected sets of instructions. The operator also may initiate changes in the 

system operating mode and enter data of his own choosing. Such keyboard entries can be viewed first on the 

plasma display panel before being entered into the 

computer through this channel.

The keyboard and plasma display combination comprises an operator terminal; the unit chosen for the 

APT system is the Interstate Electronics Corp. Model 

PD3000 (see fig. 27). The keyboard, similar in layout to 

a standard typewriter, is packaged separately for easy 

attachment to, for example, the tabletop at the APT 

system operator's position.

The ruggedized plasma display panel also is packaged separately in a metal box that is shock mounted in 

standard light-alloy racks above the keyboard at convenient eye level for the system operator. The overall 

dimensions, mass, and power requirement for the unit 

are given in table 1.

The panel display area or screen is an 8.5-in. (0.22-m) 

square grid of evenly spaced points or dots. The display 

can be operated to present point plots and graphics, and 

two sizes of internally generated alphanumeric charac-
52 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

ters are available. Dedicated memory in the unit eliminates the need for refreshing any display.

Channel number 7 (fig. 25) provides a second AS 

interface, which is held as a spare in the APT system.

Channel number 8 (fig. 25) features a Line Printer 

Controller (LPC) interface, through which an electrical 

connection is made to control the operation of a line 

printer. As part of its control function, the circuit board 

converts the selected computer output data into a format suitable for the printer.

In a generic sense, the line printer belongs with the 

other computer-oriented peripheral devices (see fig. 25) 

described in the following section. It is described here, 

however, primarily because its operational control is

through the last of the numbered I/O channels.

The printer unit selected for the APT system is the 

Miltope Corp. Model HSP3609-212A (see fig. 28), 

which is ruggedized and packaged in a metal box that 

is shock mounted in standard light-alloy racks. The 

overall dimensions, mass, and power requirement for 

the unit are given in table 1.

The printer can operate at speeds as great as 500 

lines per minute in 80-column [10 characters per inch 

(about 4 per centimeter)] or 132-column [17 characters 

per inch (about 7 per centimeter)] format on fanfold or 

friction-feed paper. The printer uses a set of 64 alphabetic, numeric, and special characters and will print, as 

programmed or commanded by the APT system operaFlGURE 27. Operator's keyboard and plasma display terminal for the aerial profiling of terrain system.
INSTRUMENT SYSTEM 53

tor, selected data and computer status reports for postflight analysis. Probably, it also will be programmed to 

print a log of all system-operator actions.

Peripheral Devices

The only remaining component directly connected to 

the UNIBUS is the magnetic tape controller (see 

fig. 25). The unit chosen for the APT system is the 

Miltope Corp. Model TC-11 (see fig. 29), which is packaged in a metal box that is mounted in standard lightalloy racks. The overall dimensions, mass, and power 

requirement for the unit are given in table 1.

As its name implies, this peripheral device controls 

the flow of data (bidirectional) between the magnetic

tape recorders and the computer. In performing the 

recording aspect of this function, the magnetic tape 

controller starts and stops the tape recorder and reformats the data in a way suitable for entry on the tape. 

When the end of a tape is reached, the controller automatically diverts the stream of data to the second 

recorder so that no information is lost. If the second 

recorder is not available, then the controller stops the 

data flow and signals the system operator with a warning light. The controller can accommodate tape speeds 

as fast as 130 in./s. At 800 bytes per inch, this represents a data rate of 104,000 bits per second.

The observations on controller capability are equally 

valid if a recorder is loaded with a magnetic tape cartridge containing data to be read into the computer. In

FIGURE 28. Line printer for aerial profiling of terrain system.
54 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

FIGURE 29. Magnetic tape controller for the aerial profiling of terrain system.

fact, in the normal operation of the APT system, the 

magnetic tape devices will be used, not only to record 

data output from the computer (and the APT subsystems) during preflight calibration and alinement procedures and during surveying flights, but also to load 

data and programs into computer memory before a 

flight mission.

The magnetic tape recorder selected for the APT system is the Miltope Corp. Model CR600 (see fig. 30), 

which is a ruggedized cartridge-load type. The recorder 

is packaged in a metal box mounted in standard lightalloy racks. The overall dimensions, mass, and power 

requirement are given in table 1. Note (fig. 25) that two 

tape recorders will be used to forestall any break in 

recording data when a full cartridge must be replaced.

A cartridge contains a standard IBM-compatible 7- 

in. (0.18-m) diameter reel capable of storing 600 ft 

(183 m) of V2-in. (13-mm)-wide magnetic tape. Data are 

stored on the tape in a seven-track recording mode at a 

density of 800 bits per inch (315 bits per centimeter). 

The tape recording speed is 25 in./s (0.64 m/s) and 

rewind speed is 50 in./s (1.27 m/s). These specifics indicate that a single cartridge (tape) can hold 6.24 million 

bits of data and (or) information, which roughly translates (not a simple or direct conversion) into about a 

30-minute surveying flight.

The line printer, which is the only remaining 

computer-oriented peripheral device shown in figure 25, has been described in the discussion on I/O

channel number 8, given in the section "Subsystems 

and Data Interfaces."

Software

Software is, for this report, the ensemble (regardless 

of format) of instructions, commands, and programs 

needed to run the computer through all its necessary 

functions in all desired modes of APT system operation. 

A program is defined as a set of instructions or steps 

that tells the computer exactly and completely how to 

handle a given task.

If the software for the APT system is to fulfill its 

basic purpose, it must be keyed to a very rigorous 

mathematical analysis and description of how that system performs. Among the powerful analytical tools 

used in the mathematics that describe the APT system 

is the method of "state variables." This method is 

ideally suited for analysis of the multiple-input and 

multiple-output APT system, where performance details can ony be described completely if the values of, 

and changes in, more than 50 variables can be determined. In applying the method, it was used in a nondeterministic way; that is, it was used to get "best estimates" which closely resemble what might be obtained 

through the analytical technique of least-square fits. 

The end results of such a sophisticated mathematical 

analysis and the processing of such large bulks of sys-
INSTRUMENT SYSTEM 55

tern operating data are judged to contribute improvement in the position-coordinate computations of about 

two orders of magnitude.

For the APT system, a unique computer program has 

been written for each of eight modes of system operation. Parts of these programs, called subroutines, are 

common to several or all system modes. Examples include instructions for the computer to perform computations, to interpolate data, to generate mathematical 

functions, to accept data (input) for storage or processing or both, and to release data (output), all as required 

by the APT subsystems and peripheral devices described in the preceding sections.

Wherever possible, the programs have been designed 

to be "table driven." This is a technique whereby essential pieces of APT system control information are represented as data items or variables, organized into tables, 

and stored in the computer memory, rather than being 

built into the program structure itself. An example is 

the requirement for periodic computation of temperatures from thermistor outputs at dozens of widely scattered points in the APT system. The instructions for 

checking individual thermistors are similar but not 

identical. Thus, to write dozens of not-quite-identical

FIGURE 30. Magnetic tape recorder for the aerial profiling of terrain system.
INSTRUMENT SYSTEM FOR AERIAL SURVEYING

instruction sets is to invite many opportunities for errors, in addition to using up too much precious memory 

space. By organizing the key control information, in 

this case, scale factors, for all thermistors into a table 

(one numbered line per thermistor) then, only one set of 

instructions is needed. For a given thermistor, those 

instructions simply reference the appropriate line 

number in the table where all the pertinent control 

information resides. Thus, any subsequent requirement for making a change in that information is easily 

met without rewriting the master program.

The abovementioned computer programs, instructions, and related earth science and instrumentcomponent-performance data needed to operate the 

APT system in all its modes are hereby labeled the 

"master operating program." This program is written 

on one magnetic tape.

The fact that the Norden computer must function in 

a number of ways as a real-time control device for the 

APT system means that a premium is placed on 

scheduling and timing the various program tasks the 

computer must execute. The message-priority features 

already built into the computer software greatly simplify the writing of programs to cover all tasks, but the 

challenge is persistent and very real to achieve maximum program efficiency so that the computer has 

"free" time between major program task cycles wherein 

the lower priority and miscellaneous program tasks 

can be fitted. Thus, the design and instruction coding of 

a particular program must reflect special attention to 

the completion of the desired work within the available 

computer time. Furthermore, the steps to be performed 

must be expressed as concisely as possible, so that the 

whole program will fit within the limited computer 

memory space.

Organization of the software for the APT system is 

better understood if prefaced by brief descriptions of 

the eight distinct modes in which the system will be 

operated. These modes encompass preflight, flight, and 

postflight operations and are identified by name as follows:

1. COMPUTER OFF.

2. STARTUP AND TEST.

3. CALIBRATE AND ALINE

4. STANDBY.

5. EN ROUTE. "|

6. SEARCH. ,. , , ,. 7. AERIAL CALIBRATE. f Fllght °Peratlons

8. AERIAL SURVEY. J

Once the APT system is operational, a typical flight 

mission might logically require it to be sequenced 

through the foregoing modes in the manner shown in 

figure 31. Frequent reference to that figure will give 

added meaning to the ensuing mode descriptions. 

1. COMPUTER OFF. Mode in which no electrical

power is fed to any part of the computer or its peripheral devices. The mode is characterized by a preestablished arrangement of switch settings and specified 

values for certain internal state-of-health sensors. In 

this mode, it is possible to run all analog circuitry 

which includes servomotors and temperature controls. 

Thus, the IMU can be powered and gimbal angles set.

2. STARTUP AND TEST. Mode used to power up 

the system, perform functional system testing, and enter data pertinent to operation in flight. The computer, 

peripheral devices, and APT subsystems are powered 

up through a preestablished sequence of steps, which 

are executed by manual reference to a checklist and 

subsequently by operator selection on the keyboard. A 

key step in this sequence includes insertion into a taperecorder unit of the magnetic tape cartridge that contains the "master operating program." This is read into 

computer memory by means of hard-wired instructions 

invoked by operator selection on the computer control 

panel. When a signal shows the program to be properly 

loaded in the computer memory, the "executive program" is called up, and the first instructions therein 

are executed when the computer receives its next 200- 

Hz time signal from the master clock. The significance 

of this and other time signals, and a brief definition of 

the executive program, are described later in this section.

Because the keyboard-and-plasma-display terminal 

is the principal interface between the operator and the 

APT system and provides the means for shifting the 

system from one operating mode to another, its description needs some amplification at this point. The plasma 

display is divided operationally into two parts, a soAERIAL 

CALIBRATE SEARCH

FIGURE 31. Hypothetical sequence of the aerial profiling of terrain system modes for a flight mission.
INSTRUMENT SYSTEM 57

called "fixed" upper segment and an "interactive" 

lower segment.

In a preestablished sequence and format, the fixed 

segment displays information relevant to the current 

operating mode and its status. This will include, for 

example, current readings of the IMU gimbal angles 

and similar current real data for the particular components of the APT system that are being functionally 

tested at that moment. This upper part of the display is 

updated by the computer every 2 seconds.

The interactive lower segment displays a "menu" of 

numbered alteratives from which the operator may select by means of the keyboard. Among the alternatives, 

for example, will be certain component test routines 

that can be started, other information displays that can 

be called up, or opportunities to enter data into the 

computer or to shift the APT system into another operating mode. The operator makes a selection through 

the keyboard, and the computer takes the commanded 

action when the carriage return key is depressed.

No attempt is made here to detail all I/O actions 

involving the computer, inasmuch as most will appear 

in a subsequent table and several figures. It is sufficient simply to observe that the STARTUP AND TEST 

mode provides for the orderly and safe automatic 

startup (or shutdown) of the APT system. The procedures include certain computer self-checks, various 

subsystem tests, and continuous alarm monitoring 

keyed to preestablished limits for automatically read 

voltage and temperature sensors at critical places in 

the system.

3. CALIBRATE AND ALINE. Mode used to finetune the inertial navigator while the aircraft remains 

standing over a known point on the airport parking 

ramp. Once the operator has selected this mode and 

commanded the computer through the previously described keyboard actions, a preestablished sequence of 

procedures is started. These include the deliberate precessing of the stable platform in the IMU through a 

prescribed schedule of attitudes by appropriately 

torquing the gyros. The need is to determine (calibrate) 

all gyro and acelerometer performance coefficients at 

this particular point in time and under the prevailing 

set of environmental conditions which include the 

physical location of the APT system on the Earth's surface.

The calibration exercise is an estimation process, 

often loosely called "Kalman filtering." It compares 

previously designed probability models of individual 

accelerometer and gyro performance against their actual physical measurement of the unchanging and 

known gravity and Earth-rotation vectors in the stationary aircraft. The desired performance-coefficient 

estimates emerge from the processing of the foregoing 

data over a period of from 1 to 48 hours, depending on

the precision level required. The ramifications of the 

Kalman filtering technique for the APT system will 

become more apparent in the section on "Postflight 

Data Processing." However, the following general definition of a Kalman filter is offered as a useful preface. 

The Kalman filter is a mathematical procedure (or 

algorithm) for combining measurement data with information about the internal configuration and performance characteristics of a system to form estimates 

of the current state of the system. It is named after its 

inventor, R. E. Kalman (1960), and is called a filter, 

following well-established usage, because its effect is to 

"purify" error-contaminated, or "dirty," data signals, 

thereby allowing the system state to be observed. Although the Kalman filtering procedure is conceptually 

akin to taking the average of replicate measurements, 

or fitting a regression line to a scatter-plot, the following considerations distinguish it from such elementary 

statistical procedures:

1. The state of the system is not directly observable but 

must be determined by calculation from measurable data.

2. The measured data are subject to unavoidable measurement errors, so the states calculated from 

them also are subject to error.

3. The true state of the system is changing at roughly 

the same rate as the rate at which data are being 

acquired, so that replicate observations are not 

available for averaging.

4. The complexity and volume of the necessary computations tax the ability of the computer to produce 

state estimates as fast as they are needed. 

To cope with these problems, the Kalman filter is recursive it computes its state estimate in the form of a 

weighted average of the previously computed estimate 

and an estimate based on the most recent set of measured variables. In addition, the filter maintains a running estimate of the statistical sampling errors in these 

estimates. In this way, it can compute the optimum 

weighting factor to minimize the error of each estimate.

The calibration exercise is complete once a current 

set of performance-coefficient estimates is available for 

all gyros and accelerometers, which the operator considers adequate to predict the performance of those instruments in the upcoming APT system flight modes. 

At this point in time, alinement begins, wherein the 

operator invokes computer functions that direct the 

procedures for stable-platform leveling and gyrocompassing (see "Performance Dynamics" section). The 

minimum time required for the alinement exercise is 

about 20 minutes.

The APT system operator may selectively opt to tape 

record or print out data produced during the calibration 

and alinement exercises. Normally, the operator will
58 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

allow the system to continue its alinement routine 

until shortly before flight time. At that point the operator will command the system to shift to the 

STANDBY mode (see fig. 31).

4. STANDBY. Mode in which the APT system is 

maintained in powered up, calibrated, and alined configuration in anticipation of a flight mission. This program segment includes provision for inputs to computer memory from the static-air-pressure and 

temperature transducers mounted on the outside of the 

aircraft.

While the APT system is in STANDBY status, the 

operator is responsible for keying in the pertinent 

flight-mission data. These will include, for example, 

information on the flight path to be navigated to the 

survey area, position-coordinate and gravity values for 

the field-surveyed retroreflectors (minimum of three 

required), and details of the aerial survey to be accomplished. Any last-minute changes in the flight mission 

also would be keyed in at this time.

When all preparations for the flight mission have 

been completed and the flight crew is ready, but before 

the aircraft starts to move, the system operator will 

command (through the keyboard) the system to shift to 

the EN ROUTE mode (see fig. 31).

5. EN ROUTE. Mode in which the APT system will 

measure and record the actual path of the aircraft as it 

moves away from the ground-alinement location. It is 

also the mode in which the system will navigate the 

aircraft over the flight-plan route to the survey area.

When the system operator has commanded this 

mode, the computer functions in this program segment 

include processing of the measurement data from the 

three accelerometers and from the static-air-pressure 

transducer to accomplish inertial navigation and computation of steering signal outputs for the indicator on 

the pilot's instrument panel so the intended flight path 

can be flown more precisely.

Although tacitly implied, it bears repeating that the 

computer functions in this mode, as in all modes, continue to include the monitoring of state-of-health sensors throughout the APT system. Status readings appear at preprogrammed intervals on the operator's 

plasma-display terminal. A suitable alarm will be signaled if any reading is out of bounds.

At predetermined regular intervals in the EN 

ROUTE mode, the computer will route to the plasmadisplay terminal instantaneous readings of the aircraft 

position coordinates, heading, ground speed, and the 

time and distance remaining to the first presurveyed 

retroreflector target that is to be used in the survey. 

The system operator also may opt to tape record or 

print out any EN ROUTE data considered appropriate 

for use in the postflight data-processing work.

As the aircraft nears the first retroreflector, the system operator will command the system to shift to the 

SEARCH mode (see fig. 31). If the circumstances warrant, however, the operator may choose to command a 

shift to any other flight mode.

6. SEARCH. Mode in which the tracker will be run 

through its search pattern until the first presurveyed 

retroreflector target is found. When the APT system 

operator has commanded this mode, the computer functions now include, in addition to those that concern 

continued EN ROUTE navigation and steering, command of tracker gimbal torques sufficient to point the 

tracker forward toward the expected location of the 

retroreflector and to begin automatically to scan 

around the pointed direction when the retroreflector 

comes within range, with the intent of detecting and 

locking onto a pulsed-laser return signal.

The nature of the program segment, written to cover 

this computer role of commanding the tracker in its 

search patterns, can be understood through a brief account of what the computer does. The computer first 

selects the nearest retroreflector and computes its 

range. If it decides the target is out of range, then the 

computer simply commands the tracker to point forward at the maximum angle. If the target is determined to be in range, then the computer estimates the 

probable error in target position. If the estimated error 

is 65 ft (20 m) or less, then the computer commands the 

tracker to begin a spiral scan that gradually opens up 

around the initial most probable direction. The computer examines successive groups of return signals 

from the tracker quadrant photodiode and range detectors, and, when certain reasonableness criteria are 

met, it concludes that a valid return signal has been 

received. At this point, the computer commands the 

tracker to enter its tracking procedure, which means 

that the gimbal-drive servos, using the quadrant 

photodiode signals, maintain the pointing direction 

needed to keep the tracker locked onto the target. The 

computer simultaneously and automatically shifts the 

APT system to the AERIAL CALIBRATE mode (see 

fig. 31).

If, in the foregoing procedure, a valid return is not 

received within 15 seconds, then the computer terminates the spiral scan and begins a wide-angle, side-toside, or raster, scan, around the most probable position 

of the retroreflector. Upon receipt of a return signal 

within the range-gate setting chosen for this fast raster 

scan, the computer points the tracker in the direction of 

the valid return and once again commands the spiral 

scan. The lock-on procedure is as described above.

If the error in retroreflector position, as first estimated by the computer, is greater than 65 ft (20 m), 

then the computer simply commands the fast raster
INSTRUMENT SYSTEM 59

scan at the outset, and the ensuing procedures to target 

lock-on are as described above.

7. AERIAL CALIBRATE. Mode in which the 

tracker individually tracks, according to the sequence 

stipulated in the flight plan, each retroreflector in the 

survey area. Whenever the computer has initiated this 

mode automatically, the pointing-error signals from 

the quadrant photodiode replace the computergenerated, angle-command signals feeding the tracker 

gimbal-drive servos, target lock-on is achieved, and 

tracking the retroreflector begins. Other computer 

functions continue as in the EN ROUTE mode. On any 

interruption of tracker returns (caused perhaps by obscuration of the laser-beam path to the retroreflector), 

as signalled to the computer by the tracker circuitry, 

the computer will command reentry to the SEARCH 

mode to reacquire the retroreflector and then continue 

tracking. As the tracking procedure progresses, a 

stream of readings of laser range, tracker pointing angles, and IMU gimbal angles accumulates in the computer storage. From these data, the computer refines 

the indicated aircraft position and computes preliminary position data for the retroreflectors.

The refinement process is aided by computer corrections for the tracker measurements, based on index-ofrefraction calculations for the air path traveled by the 

laser beam. The adjusted tracker measurements are 

formatted and recorded for the postflight dataprocessing work. This work will include adjustments 

on the whole body of flight-mission data so that the best 

possible final values are determined for such key 

parameters as position coordinates for each unsurveyed retroreflector and for the aircraft at the succession of points that defines its flight path.

If the tracker looses lock on a retroreflector, then the 

computer will determine whether that target is still 

within range. An affirmative result will trigger automatically a slow-scan reacquisition procedure, inasmuch as the target location is already rather precisely 

known. When the tracking of a particular retroreflector is complete, as shown by loss of lock and failure of 

the range check, the computer determines the heading 

to the next retroreflector, and this result is relayed to 

the pilot's steering indicator and to the APT system 

operator's plasma display.

Normally, the AERIAL CALIBRATE exercise is 

complete when each of three specified retroreflectors 

has been individually tracked twice in the carefully 

designed flight-plan sequence. If necessary, however, 

the system operator can alter the preplanned exercise 

by adding or deleting certain tracking passes.

When the operator is satisfied that all retroreflector 

targets needed for the flight mission have been properly tracked, the sytem is commanded to shift to the 

AERIAL SURVEY mode (see fig. 31).

8. AERIAL SURVEY. Mode in which the APT system collects the survey data for which the flight mission was planned. When the system operator has commanded this mode, the computer functions in this 

program segment include, in addition to those that concern continued inertial navigation and steering, the 

slow-scan search and reacquisition and tracking procedures for each retroreflector that is to be used during 

the actual gathering of the aerial survey data. The 

sequence is as stipulated in the flight plan for this part 

of the mission. These functions draw upon the inputs 

made by the system operator during the STANDBY 

mode.

After commanding entry into the AERIAL SURVEY 

mode, the system operator will command activation of 

the laser profiler and color TV camera subsystems. 

Computer functions then expand to include performance monitoring of these subsystems, relaying the APT 

system time-code signals to the time-date generator 

(see "Color TV Camera" section), relaying the timelabeled video imagery to the video recorder, and formatting and tape recording the terrain-profile data 

that emerge from the laser-profiler range measurements. As recorded in flight, the position coordinates 

for the profile data are in a raw state; refinement will 

come through the postflight data-processing work.

When all the desired terrain-profile data have been 

gathered, the operator may command the system to 

shift back to the AERIAL CALIBRATE mode to allow 

the collection of a final suite of position data for the 

retroreflectors used in the mission. The operator then 

may command the system to shift to the EN ROUTE 

mode to navigate to the next survey site or to return to 

the airport of origin.

Tacit in the foregoing mode descriptions is a natural 

and orderly progression from one to the next, with opportunities to jump back one or more modes as operating procedures dictate.

In the foregoing decriptions of the eight distinct APT 

system modes, many allusions have been made to the 

various computer functions that make the sytem perform in the desired manner. Those functions are performed successfully, however, only as a consequence of 

the sophisticated structure and organization of the software. The challenge now is to give some meaningful 

description to that structure and organization. As this 

descriptive material is read, it is important to remember that all programs developed for the APT system 

were designed to run under the Digital Equipment Corporation's (DEC) POP 11/70 "operating system program" featuring the Resource Sharing Executive Program, RSX-11S.

The primary purpose of the operating system program is to manage the execution of all program segments coded specifically for the APT system. For each
60 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

APT system mode, proper management requires that 

the program segments, hereafter called tasks, be run in 

the correct sequential relations with respect to timing 

and priority. A function of the RSX-11S, which is incorporated in the operating system program, is to establish a priority queue for the tasks that must be run 

and to handle a task interrupt when a higher priority 

need arises. An interrupt entails temporary storage of 

the uncompleted tasks, together with all variables and 

intermediate results, until it can be resumed when the 

higher priority work is done. In reality, several layers 

of interrupted tasks can be handled as the need arises.

The operating system program also includes I/O 

driver routines (subroutines that feed data into or out 

of peripherals when called), a library of mathematical 

and other functions used in running many tasks, and 

definition of common storage areas (particular portions 

of computer memory used for communication among 

various tasks).

The first noteworthy developmental milestone was 

reached when the PDF 11/70 operating system program, generated for the APT system software configuration, was run in the Norden computer. This milestone program served as the framework around which 

the balance of the program development was fitted. As 

the development work proceeded, program segments 

were written or coded in either Fortran or Macro-11 (a 

DEC language) assembly languages. The segments 

were checked by test runs on the Draper Laboratory's 

VAX 11/780 computer (a DEC machine), which can be 

configured to emulate the Norden POP 11/70M computer. The VAX computer then produced the machinelanguage version of the master operating program that 

is loaded into the Norden.

In addition to the operating system program described above, the master operating program includes 

the 14 special APT system tasks that are used in various combinations to meet the performance requirements of each of the eight operating modes. The structure graph in figure 32 shows the component parts of 

the operating system program.

Note in figure 32 that the 14 special tasks are organized logically under six basic operational functions 

within the master operating program. Of the six functions, only the first, processing that must be done in 

real time, contains tasks that are to be run at stipulated frequencies, which are submultiples of the APT 

system master-clock frequency. Because of this linkage 

to the master frequency, the five tasks shown (fig. 32) 

in this category are termed "synchronous tasks"; the 

first four are named by their recurrrence intervals of 

execution, and all five are listed in their proper activation (callup) sequence.

The remaining five basic operational functions embrace the nine remaining special APT system tasks,

and, because each one can be called by the APT system 

operator or by another task to run at any time, they are 

all termed "asynchronous tasks" (see fig. 32). All nine 

are lower in priority than the five synchronous tasks, 

and, hence, when one is called, it must be fitted into the 

priority queue (by the executive program) and await 

available time on the central processor.

Although the labels for the foregoing five basic operational functions (fig. 32) are reasonably self-evident, 

some amplification is appropriate. "Hardware mode 

control" is the single all-important asynchronous task 

of monitoring and controlling the shifting of the APT 

system from one mode to another. "Telemetry and 

alarm processing" are two asynchronous tasks that 

concern the monitoring of all system-health sensors, 

logging their readings, and alerting the system operator to any malfunctions. "Data base population and 

recording" are also two asynchronous tasks that allow 

the operator to bring in (to computer storage) programs 

and data preassembled on magnetic tape or permit 

manual entries by means of the computer keyboard and 

to unload (onto magnetic tape), during and after a 

flight mission, the complete suite of data used in the 

mission, as well as the survey data collected. "Hard 

copy command processing" includes three asynchronous tasks that relate exclusively to calls for output on the line printer. "Operator system interface" is 

the single and last asynchronous task that governs the 

APT system operator's keyboard and plasma-display 

terminal.

Each of the 14 special tasks (program segments) 

shown in figure 32 embraces a modest variety of computer processing actions. These are given in summary 

form in table 2, which also shows how each special task 

is activated (called) and the particular APT system operating mode(s) in which each computer processing action occurs. The table is quite complete and displays, to 

the intended degree, the range in complexity of the 

computer programs developed to operate the APT system. A more complete description of these programs is 

beyond the scope of this paper.

Just a cursory survey of table 2 shows that most of 

the computer processing actions relate to the five realtime synchronous tasks. Furthermore, for a given synchronous task, the variety of computer actions listed is 

to be recognized as a simple collection of those actions 

that have the same priority and required frequency (or 

frequency submultiple) as the task itself. Those actions 

that need to be accomplished at the same frequency as 

the task are grouped under the heading "Primary"; 

those actions that need only be accomplished at some 

specified submultiple of the task frequency are grouped 

under the heading "Secondary."

In the group of five synchronous tasks, the highest 

priority and frequency (shortest recurrence interval) 

are assigned to the TWENTY MS task. The other four
INSTRUMENT SYSTEM 61

EXPLANATION

(3) PRIORITY LEVEL

OPERATING SYSTEM 

PROGRAM

(EXECUTIVE,

I/O DRIVERS,

LIBRARY ROUTINES,

COMMON STORAGE AREAS)

FOURTEEN

REAL-TIME 

PROCESSING

>i

"-o

£1) TWENTY MS 

J2) EIGHTY MS

(3) TWO SECOND 

J5) SIXTY SECOND

(4) CALIBRATE AND 

~~ ALINE

SPECIAL TASKS

TASK CATEGORY

\

TELEMETRY 

AND ALARM 

PROCESSING

DATA BASE

POPULATION

AND RECORDING

OPERATOR

SYSTEM 

INTERFACE

8) MODE TRANSITION (10) TELEMETRY

(8) ALARM MONITOR

(7) DATA INITIALIZATION 

(6) DATA RECORDING

(81 ERROR MESSAGE 

PRINTER

(9) LIST PROCESSOR

(8) PRINTER CONTROLLER

1 

|_(8(8) DISPLAY

NINE ASYNCHRONOUS TASKS 

FIGURE 32. Structure graph of the master operating program for the aerial profiling of terrain system.

tasks are assigned lower priorities and frequencies, but 

each of the latter is a specified submultiple of the 

highest frequency. This leads to a simple callup sequence for the five synchronous tasks, and a logical 

grouping of the primary and secondary computer processing actions thereunder, as outlined in figure 33. In 

effect, this illustration serves to emphasize the sequential relations that are more subtly discernible in 

table 2.

From the information given in figures 32 and 33 and 

table 2, a general statement naturally evolves to summarize the callup sequence for the 14 special tasks in 

the APT system. The computer processing actions proceed from the highest to the lowest priority tasks, 

which mean that all five synchronous tasks are completed before the nine asynchronous tasks. The latter 

then are completed in accord with their individually 

assigned priorities. The entire train of events is initiated by a 200-Hz (5-ms-period) signal from the master 

clock, which invokes the transfer of measurement data 

between the APT subsystems and the computer by 

means of the DMA channel (number 1 in fig. 25). When 

all system tasks are up-to-date, the computer is idle 

momentarily, awaiting the next 200-Hz signal.

The stage has now been set to conclude the discussion 

of "Software" with three functional diagrams (figs. 34- 

36) to illustrate how tasks are linked and how the data 

flow to bring about the computer processing actions 

needed for each operating mode of the APT system. In

the three diagrams, the conventions used are to show a 

given computer processing action as a rectangular box 

in which an abbreviated action title appears. The subtitle at the base of the box is generally the special task 

under which the action occurs; occasionally, it will be a 

subroutine of the master operating program, as, for 

example, the DMA I/O DRIVER. Note that color coding 

identifies computer actions with the particular APT 

subsystem(s) for which they are being performed. Thus, 

orange indicates actions related to the IMU and inertial navigation, green relates to the tracker, and blue 

relates to the profiler. Note also that, in each diagram, 

the data flow and the action proceed generally from left 

to right. Three initiating or input sources, namely, 

DMA data transfer, operator's keyboard, and analog 

telemetry, are on the left; outputs to the APT subsystems and peripheral devices are near the right. Lines 

that connect boxes signify the transfer of data or commands. A dot where lines cross indicates a physical 

connection; absence of a dot indicates no connection.

The computer processing actions in the STARTUP 

AND TEST mode (fig. 34) use all input and output 

channels of the computer in powering up the system, 

testing for proper functioning, and entering basic data. 

Many of the computer actions involve formatting data 

and relaying information through channels used for 

more complex and sophisticated functions in subsequent operating modes.

The computer processing actions for the CALI-
62 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

TABLE 2. Computer-action summary for the 14 special tasks in the aerial profiling of terrain system

TASK NAME ACTIVATION COMPUTER PROCESSING ACTION APT-SYSTEM MODES*

TWENTY MS Called by the direct 

memory access (DMA) 

input/output (I/O) 

driver every fourth 

DMA data transfer.

Primary:

Compute IMU and tracker gimbal angles and accelerometer 

head angles from DMA data transfer, and sum for 80 ms 

averaging.

Compute IMU and tracker gimbal rates. 

Detect IMU and tracker gimbal movement.

Sum and save profiler and tracker time-interval data

from each DMA data transfer for computation of laser ranges.

Compute IMU and tracker gimbal-drive signals.

Compute gyro torque pulse sequences to be applied each 

5 ms over next 20 ms.

Digital-to-analog output signal processing. 

Save data for recording.

Secondary:

Forty ms (Every other pass):

Average profiler time-interval data for later computation of 

laser ranges.

Save data for recording.

Eighty ms (Every fourth pass):

Reorder and transfer partially processed tracker-range data 

for EIGHTY MS task.

Save gimbal and accelerometer head-angle data for EIGHTY MS 

task.

Call EIGHTY MS task. 

Save data for recording.

Two Second (Every 100th pass): 

Call TWO SECOND task.

Save data for recording.

Sixty Second (Every 3000th pass):

Sample and save accelerometer data for validation in 

SIXTY SECOND task.

Call SIXTY SECOND task. 

Save data for recording.

2-8 incl.

2-8 incl. 

2, 6, 7, 8 

2, 6, 7, 8

2, 6, 7, 8 

5, 6, 7, 8

2, 5, 6, 7, 

2-8 incl.

2, 6, 7, 8 

2-8 incl.

2-8 incl. 

2-8 incl.

2-8 incl. 

2-8 incl.

2-8 incl.

2-8 incl. 

2-8 incl.

*Explanation of code numbers:

2. STARTUP AND TEST.

3. CALIBRATE AND ALINE.

4. STANDBY.

5. EN ROUTE.

6. SEARCH.

7. AERIAL CALIBRATE.

8. AERIAL SURVEY.
INSTRUMENT SYSTEM

TABLE 2. Computer-action summary for the 14 special tasks in the aerial profiling of terrain system Continued

63

TASK NAME ACTIVATION COMPUTER PROCESSING ACTION APT-SYSTEM MODES*

DATA 

INITIALIZATION

Called by DISPLAY 

task.

Load (record) predesignated segments of data base from 

(to) magnetic tape, on a single call by DISPLAY task.

Load data manually entered by DISPLAY task into 

predesignated segments of data base.

Call and transfer predesignated segments of data base 

to PRINTER CONTROLLER task for listing, on a single 

call by DISPLAY task.

2-8 incl.

2-8 incl.

2-8 incl.

DATA 

RECORDING

Called by TWO SECOND, 

SIXTY SECOND, and 

DISPLAY tasks

Buffer data saved for recording by synchronous tasks.

Enable/disable predetermined recording lists, on 

call by DISPLAY task.

Format and transfer data from enabled recording lists 

to magnetic-tape recorder

Switch to alternate tape unit, on either tape-error 

or end-of-tape signal, without loss of data.

Transfer status of tape units, on call by DISPLAY task.

2-8 incl. 

2-8 incl.

2-8 incl.

2-8 incl.

ERROR 

MESSAGE

Called by ALARM 

MONITOR or any 

other task.

Format all system status and error information.

On call, transfer message, date, time, and identity of 

calling task to PRINTER CONTROLLER task for printing.

2-8 incl. 

2-8 incl.

PRINTER 

CONTROLLER

Called by ERROR 

MESSAGE and LIST 

PROCESSOR tasks.

Control transfer of information to the line printer on 

a "time available", non-interference basis.

Monitor printer off-line status.

Transfer status information for display every 15 

seconds, if enabled by DISPLAY task.

2-8 incl.

2-8 incl.

LIST 

PROCESSOR

Called by any processing (synchronous or 

CAL & ALINE) task or 

DISPLAY task.

On a single call by a processing task, format requested 

list from APT system data base and call PRINTER CONTROLLER 

(or DISPLAY) task to print out (or display) list.

On a single call by DISPLAY task, format and call PRINTER 

CONTROLLER (or DISPLAY) task to print out (or display) list.

2-8 incl.

2-8 incl.

DISPLAY Called by operator. Format and route operator commands to initiate, control, 

and monitor APT system moding.

Format and route operator commands to DATA RECORDING, 

DATA INITIALIZATION, and LIST PROCESSOR tasks.

Format and transfer data entered by operator. 

Format and display data requested by operator.

2-8 incl. 

2-8 incl.

2-8 incl. 

2-8 incl.

*Explanation of code numbers.

2. STARTUP AND TEST.

3. CALIBRATE AND ALINE.

4. STANDBY

5. EN ROUTE.

6. SEARCH

7. AERIAL CALIBRATE.

8. AERIAL SURVEY
64 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

TABLE 2. Computer-action summary for the 14 special tasks in the aerial profiling of terrain system Continued

TASK NAME ACTIVATION COMPUTER PROCESSING ACTION APT-SYSTEM MODES*

SIXTY 

SECOND

Called by TWENTY MS 

task.

Transform sixty-second accelerometer data to locallylevel coordinate frame.

Compute sixty-second gyro torque pulse sequences. 

Validate acceleroraeter head-angle data. 

Compute gyro-torque-test commands.

Initialize and start accelerometer coning misalinement 

estimate computations.

Compute coefficients for accelerometer coning compensation. 

Perform accelerometer diagnostic calibration. 

Call CAL & ALINE task. 

Save data for recording. 

Call DATA RECORDING task.

2, 3

2, 3, A

3

2

2

2

2

3

2, 3, 4

2, 3, 4

CAL & ALINE Called by SIXTY 

SECOND task.

Compute gyro torque commands to rotate IMU stable platform 

through a series of predetermined trajectories.

Compute estimates of IMU-performance parameters in 57-state 

Kalman filter.

Save updated estimates of IMU-performance parameters.

MODE 

TRANSITION

Called by DISPLAY, 

ALARM MONITOR,or 

EIGHTY MS task.

Monitor and control moding of IMU, tracker, and profiler. 2-8 incl.

TELEMETRY Called by executive 

program (about once 

per second).

Sample analog-to-digital converter outputs.

Smooth those outputs that indicate temperature

Convert sampled and smoothed outputs to engineering units.

2-8 incl. 

2-8 incl. 

2-8 incl.

ALARM 

MONITOR

Called by executive 

program every minute 

or by DISPLAY task.

Primary:

Monitor telemetry data, flagging and saving quantities 

crossing predetermined thresholds.

Call MODE TRANSITION task for appropriate emergency action 

according to predetermined list of quantities if enabled 

by DISPLAY task.

Call ERROR MESSAGE task to report quantities when flagged.

When called by DISPLAY task, call ERROR MESSAGE task to 

report number and status of flagged quantities.

Secondary:

Sixty Minute (Every 60th time through):

Call ERROR MESSAGE task to report number of flagged 

quantities.

2-8 inel. 

2-8 incl.

2-8 incl. 

2-8 incl.

2-8 incl.

*Exf»lanation of eode numbers:

2. STARTUP AND TEST,

3. CALIBRATE AND ALINE.

4. STANDBY.

5. m ROUTE.

6. SEARCH.

7. AERIAL CALIBRATE.

8. AERIAL SURVEY.
INSTRUMENT SYSTEM 

TABLE 2. Computer-action summary for the 14 special tasks in the aerial profiling of terrain system Continued

65

TASK NAME ACTIVATION COMPUTER PROCESSING ACTION APT-SYSTEM MODES*

EIGHTY MS Called by TWENTY MS 

task.

Primary:

Compute accelerometer head-angle averages and rates.

Compute IMU and tracker gimbal-angle averages. 

Compute tracker range for the previous 80 ms. 

Compute accelerometer coning misalinement estimates. 

Compute accelerometer coning misalinement compensations.

Transform accelerometer data to locally-level coordinate 

frame.

Compute position and velocity (inertial navigation). 

Tracker search commanding and moding. 

Compute gyro torque commands. 

Save data for recording.

Secondary:

Two Second (Every 25th pass): 

Save data for recording.

Sixty Second (Every 750th pass):

Save IMU gimbal-angle averages for gyro torque test.

Save data for recording.

2-8 incl.

2-8 incl.

2, 6, 7,

2

2-8 incl.

5, 6, 7,

5, 6, 7, 

2, 6, 7, 

5, 6, 7, 

2-8 incl.

2-8 incl.

TWO SECOND Called by TWENTY MS 

task.

Compute IMU and tracker gimbal-angle commands.

Compute position and velocity update.

Compute heading vectors and time to reach retroreflectors.

Compute flight-indicator signals.

Format APTS date and time for displays, printer, and recorder,

Save data for recording.

Call DATA RECORDING task.

2

7, 8

5, 6, 7, 

5, 6, 7, 

2-8 incl. 

2-8 incl. 

2-8 incl.

*Explanation of code numbers:

2. STARTUP AND TEST.

3. CALIBRATE AND ALINE.

4. STANDBY.

5. EN ROUTE.

6. SEARCH.

7. AERIAL CALIBRATE.

8. AERIAL SURVEY.
66 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

200 Hz (5 MS) 

" MASTER TIMING SIGNAL

I- 80-MS PRIMARY ACTIONS 

I- 2-SECOND SECONDARY ACTIONS 

I 60-SECOND SECONDARY ACTIONS 

2-SECOND SECONDARY ACTIONS

2-SECOND PRIMARY ACTIONS 

 60-SECOND PRIMARY ACTIONS

80-SECON D SECON DARY ACTIONS

FIGURE 33. Recurrence intervals and callup sequence for the five 

synchronous tasks in the aerial profiling of terrain system.

BRATE AND ALINE mode (fig. 35) show the same 

basic structure and organization as for the STARTUP 

AND TEST mode, with the important addition of the 

special actions needed to calibrate and aline the APT 

system. The STANDBY mode continues the use of the 

same computer-action structure to hold the system in 

its alined configuration ready to perform a flight survey and to allow time for entering flight-mission data 

that will be used by the computer in the flight modes 

that will follow.

Computer processing actions for all four flight 

modes EN ROUTE, SEARCH, AERIAL CALIBRATE, and AERIAL SURVEY have enough similarities in basic structure and organization so that they 

are shown conveniently in a single diagram (fig. 36).

Figures 34 through 36 faithfully portray all computer processing actions given in table 2, although 

some actions have been lumped to improve clarity. It 

should be emphasized, however, that the computeraction linkages and flow of data shown in these figures 

are, to some degree, schematic, or analogous, to 

"snapshots" of many actions that are commingling in 

apparent kaleidoscopic fashion. An infinite number of 

such snapshots are possible, but the several chosen 

serve to illustrate important features of the many necessary computer actions. To delve any deeper into the 

real complexities of the software development is well 

beyond the intended scope of this paper.

Laser Profiler

As already described in tracing the history of this 

instrument development project, the beginning centered on the innocent question, "Is it possible to measure terrain profiles to a specified high degree of precision from an aircraft?" The precision requirements in 

fixing aircraft position in the vertical and horizontal 

coordinate directions were what gave pause in answering the question. An immediate "yes" seemed possible, 

insofar as the instantaneous electromagnetic measurement of distance from aircraft to terrain was concerned. 

Subsequent events showed, however, that even this 

partial answer was brash to a degree because the design of the laser profiler encountered its own set of 

unique challenges.

Measurement of distance or length is so fundamental 

a problem that much attention is given to it in most 

studies of basic physics. Earth scientists in general are 

familiar with the platinum-iridium bar, exactly 1 m in 

length, which, for many years, has been maintained 

under precisely stipulated and controlled conditions at 

the International Bureau of Weights and Measures to 

represent a standard unit of length used throughout 

most of the scientific world. In 1960, the eleventh gathering of the General Conference of Weights and Measures on the International System of Units (SI) redefined the meter in the more rigorous terms of a 

specified number of wavelengths of electromagnetic radiation corresponding to the decay of the krypton-86 

atom between two particular activity or product levels. 

Thus, in this part of the overall instrument development project, it was fundamentally sound to set out to 

measure a length with a laser beam of light.

The basic processes involved in such a measurement 

technique require a capability for timing the known 

velocity of light propagation [about 186,281 mi/s 

(299,790 km/s)]. If the length measurement is to be 

precise to ±0.5 ft (0.15 m), this means the timemeasuring capability must reach a precision of 1 ns, or 

the time required for the light beam to advance about 

1 ft. Tremendous developments in timekeeping by electronic means now make it possible to obtain "off-theshelf equipment equal to the nanosecond timing task; 

special purpose equipment can do even better.

In essence, the measurement technique entails generation of a light beam by a laser source. The light thus 

produced has some very unique and desirable properties, the ones of principal interest being the spectral 

and temporal coherence of the beam. Thus, when the 

light beam is generated, successive wave forms have 

the same frequency, phase, amplitude, and direction. 

Although the particular laser chosen for the terrainprofiling task does not have as high a degree of coherence as most other lasers, it can be packaged so compactly that its small emitting area literally can be
INSTRUMENT SYSTEM 67

be

c

TJ 

O

E 

H

C/3

w 

H

Q

O

C

rt

rt

Q
68 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

Vandal*

Accelerometer Head Angla Data

*

.

1

Compute Aec*ferain*t*r Comrto. C0fflp*n»atlortt

i

Compute Acc*Mrom*l*r Haad Ano.lt Average* And Rates

t

Compute AcceWromeier Head Angto*

SIXTY SECOND

S

1

i

i

MSTWENTY 

Output Data Tranafer

Compute Torgue OyroPataa

i

i

Castration General* Commanda. Trajectory Eethnate AndUpdate Parameters tnvrtlal 

» *

Transform Acealaromater DataTo LocaMy-Lavei Coordinate Fram*

1 __ 1

)

<

3

O hB

i

a

|

s

t

Q

1 

in

I

r

DMA I/ORIVER

SIXTY SECOND

S 

o

 

1 

1-

«

i

5 f

[gig

fill :I I 1"

rrnn" " "

liL!: ;:

j« z

if £

I

1

a>

t

Pfinlw Formal 

II 

 - S

S <B 

= a

a o

!

{?!

EL *- ii

1

si 1°

¥

DATA RECORDING

9

 1

LIST PROCESSOR

|TWO SECOND 

t

I

T 

 1

 

t1

1

* 

e 

C

|o

i T:

' 

:

-1

i

R

C 

1 

1 

I

;

b

1

.

PRINTER

'

DATA INITIALIZATION

t 

1

k

.._

1>

^ <

r 

u

j j 3 

r

E 

3 

J

ERROR MESSAGE

ALARM MONITOR

"s

s III 11£

J"|

 1

DISPLAY

O 

u

LANATION OF COLOR

! INERTIAL

U 1 

11

i 

} 

ic 

u

( 

(

c

 PROFILER

Word Processmo

1

1 

I

O

5X

E

II

MODE TRANSITION

OJ

1/3

jMODE TRANSITION filing of terra

0>-i

rt

X 

(-.

o 

rt

O

<*-, 

O

1/3

T3 

0

CQ

H

c

rt

H3

H

<

K

rr\

~^j ALL OTHER

the CALII or 

ILI z

G

rt

Q
INSTRUMENT SYSTEM 69

» k. s| i!

.,, ,

o

0

fA RECO

o

'

§K

il

5

E

PRINTE OMTBOLt

U

D

O 

bB

-o 

o

E

be 

S
70 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

placed at the focal point of the transmitting optics. 

Thus, the divergence of the transmitted beam can be 

extremely small, of the order of 1 mrad, or about 

3 minutes of arc. In contrast with more conventional 

light forms this results in very tight control of the 

"footprint" at the land-surface target.

The balance of the measurement process includes the 

deliberate labeling of selected outgoing light-wave 

groups by appropriate electronic modulation so they 

can be recognized when they return, provision of the 

aforementioned optical system to project the outgoing 

laser beam in a slightly diverging conical shape, provision of a second optical system either alongside or concentric with the first to collect and focus onto a photodiode detector part of the beam reflected back from the 

land-surface target, and suitable electronic amplification and processing of the returned light signals to provide the desired length or range readout data. The aggregation of components designed to perform these 

functions has been termed a "laser profiler" because of 

its planned use in collecting terrain profiles from an 

aircraft.

Because the laser profiler is intended as an adjunct to 

the APT system, its principal features must satisfy the 

same general body of design criteria which includes 

stipulations that limit bulk size, weight, and electrical 

power consumption. At the outset, the availability of 

commercial laser units was examined, but it was 

quickly concluded that manufacturers had done relatively little to capitalize on recent state-of-the-art 

developments capable of achieving much-desired reductions in size, weight, and power. Thus, the Survey, 

with key help from the Harry Diamond Laboratories, 

chose to develop its own design for the laser profiler 

and negotiated one phase of its contract with Draper 

Laboratory to accomplish the necessary laboratory experiments that led to the first "brassboard" working 

model. This approach served to maximize the prospects 

for well-designed interfaces with the APT system.

After investigating the advantages of new miniature 

solid-state components for the laser source, it was decided to use a development by Harry Diamond Laboratories that featured a simple pulsed GaAs-injection 

laser source. This particular semiconductor source is 

rugged, reliable, and inexpensive. It can be readily 

modulated at a very high rate, which is a key property 

in distance measurement, and, with its corollary electronic components, it offers high accuracy and nonambiguous range resolution along with small size, weight, 

and power consumption. It requires no water cooling or 

precisely controlled temperature environment.

At this point in time, the Survey and Draper Laboratory learned of successful test flights of a pulsed GaAs 

laser-ranging device independently developed for the 

Naval Research Laboratory. This was flown about 500

ft (152 m) above the sea surface to collect profile data on 

waveforms. Draper Laboratory obtained a working 

copy of the chassis, optics, mechanical components, and 

some electronic circuitry for the pulsed laser transceiver (see fig. 37) from the contractor.

Draper Laboratory's final design of the laser profiler 

reflects not only the many inputs from Harry Diamond 

Laboratories and the Survey, but also the experience 

gained from two sets of Survey-designed flight tests. 

The salient laser profiler features are summarized in 

this section. For added technical details, see especially 

the reports by The Charles Stark Draper Laboratory, 

Inc. (1977), Mamon and others (1976, 1978), and Youmans (1977).

An ultrashort-pulse laser source of GaAs permits use 

of a high-pulse repetition rate (3.2 kHz) and hence the 

integration of a large number of return pulses, which is 

an important feature when the target (terrain) changes 

rapidly with the forward movement of the aircraft. Furthermore, the pulsed laser unit, in contrast with a continuous wave unit, gives the capability for discriminating between multiple returns from the individual 

pulses when foliage covers some of the terrain. By 

providing an operational option for a last pulse selection technique in examining the return-signal data, 

the number of range determinations to the desired terrain surface can be maximized.

The laser transmitter module features a three-stack 

laser driven by an avalanche transistor in a hybrid 

circuit designed and described by Vanderwall and 

others (1974). This produces light pulses, at a wavelength in the near infrared of 904 nm, with 25-W peak 

power, subnanosecond rise times, and widths of less 

than 10 ns. Such characteristics ensure complete satisfaction of the stringent Federal safety standards (eyesafe at zero range) with ample reserve margin and 

handily set the stage for vertical range determinations, 

up to flight altitudes of about 2,500 ft (760 m), that are 

well within the stipulated precision of ±0.5 ft (0.15 m).

The chassis and mechanical and optical components 

for the transceiver unit (fig. 37) are basically those 

built by Associated Controls and Communications, Inc. 

of Lynn, Mass., for the Naval Research Laboratory. 

Side-by-side optics, compensated for spherical aberrations, are featured with a 6-in. (0.15-m) diameter transmitter lens and an 8-in. (0.20-m) receiver lens. Transmitted beam divergence is about 1 mrad, and receiver 

field of view is about 2.6 mrad. An optical filter in the 

receiver has a light transmission efficiency of 0.8 and a 

bandwidth of 25.6 nm.

A key mathematical expression used in sharpening 

some design choices for the laser transceiver is the socalled range equation, which offers a convenient, rigorous way of relating light energy transmitted to light 

energy returned. For this situation where the range
INSTRUMENT SYSTEM 71

Receiver 

objective 

lens

Optical 

filter and 

condensing 

lens

Avalanche 

photodiode 

receiver 

module

Laser beam

expander

telescope

Transmitter 

condensing 

lens

Laser diode 

transmitter 

and pulse 

generator .

FIGURE 37. Basic chassis with optics and some electronics for the laser profiler.

determination involves reflection of the laser beam 

back from an uncooperative target (that is, the land 

surface), the range equation has the form

= ps

IT.R2

(5)

where PR = power, receiver, 

Ps = power, source,

Tj = transmission coefficient, source optics, 

T2 = transmission coefficient, receiver optics, 

p = reflection coefficient, land surface, 

A = area, receiver lens, 

a = attenuation coefficient, atmospheric path,

and

R = range.

The interrelations among the parameters comprising 

the range equation are illustrated schematically in the 

three parts of figure 38. These parts, studied in sequence from left to right, show (A) the significant geometric features of the laser light path from transmitter 

(source) to land surface and back again, (B) the critical 

points in the flow of light energy from the lasing source 

to the land surface and back again, and (C) the mathematical description of the principal losses in energy 

flow which allows a rigorous statement to be written

that shows the net change in energy level between 

laser source and receiver.

A key factor in the design choices is to ensure that 

the receiver optics can "see" a target area slightly 

larger than the laser "footprint" and thereby intercept 

the maximum possible amount of reflected light. Figure 38 is schematic in the sense that, in the real profiler 

instrument the field of view is coextensive with the 

field irradiated.

Paramount among other fators that are critical in 

the design choices is the signal-to-noise ratio (SNR) in 

the photodiode amplifier at the focus of the receiver 

optics. The level of light energy returned to the receiver 

must be greater than the background energy "noise" in 

the receiver by a factor (the SNR) that increases with 

the precision desired in the range measurement. The 

dominant, hard-to-reduce source of receiver noise is 

solar radiation scattered by the atmosphere and terrain within the receiver field of view. However, design 

characteristics of the electronics circuits in the receiver 

quantitatively affect the SNR, and the precision specified for the range measurements of this profiler device 

requires the overall assemblage of design choices to 

yield an SNR factor of at least 5 (see Mamon and 

others, 1978, p. 872).
72 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

As mentioned in the laser-tracker description, both 

laser devices (tracker and profiler) are served by a laser 

signal processor and a time-interval counter. These 

components manage the precision timing features inherent in the outgoing laser pulses and in the return 

signals, which are so critical in the subsequent translations into precision ranges. As used in this profiler part 

of the APT system, the cited components select a maximum range (to represent land surface as opposed to a 

treetop) out of each of 8 successive measurement 

blocks, each of which contains 16 sequential range 

measurements. The eight maximums are then averaged to yield a single range value that is recorded. 

Thus, the operating resolution of the time-interval 

counter will here be its accuracy (0.4 ns) divided by the 

square root of the number of measurements that are 

averaged (eight), or 0.1 ns.

For an operating flight altitude of 2,000 ft (610 m) 

above the terrain, and an aircraft speed of 120 mi/hr or 

104 knots (193 km/hr), the foregoing profiler characteristics indicate that a vertical range measurement to the 

terrain will be saved and recorded at 7-ft (2.13-m) intervals along the flight path. Each such recorded measurement is the average of eight range measurements 

(maximums selected from 128 successive measurements) spaced along the centerline of an imaginary 

"illuminated" rectangle on the terrain about 2 ft (0.61 

m) wide and 7 ft (2.13 m) long. The horizontal position 

coordinates assigned to this recorded mesurement are 

taken arbitrarily as the center of that imaginary 

rectangle.

The art of measuring the round-trip traveltime of the 

laser profiler pulse is electronically refined yet another 

step, beyond the constant-fraction-discriminator feaLIGHT PATH IN SPACE 

(GEOMETRY)

B

FLOW OF LIGHT ENERGY 

(FIRST LAW, THERMODYNAMICS)

MATHEMATICAL 

DESCRIPTION

RECEIVER LENS 

^TRANSMITTER LENS

LAND SURFACE -*-

<A2 >

OUTGOING PULSE

MONITOR DISPLAY 

RETURN PULSE

LIGHT INTENSITY 

VECTOR (VARIES 

AS cos 6)

SURFACE

RANGE EQUATION

T-i T2CAe- 2oR "R^s-L-^ 

(SEE EQUATION 5 IN TEXT)

LIGHT REFLECTION 

(LAMBERT'S LAW)

TIME

FIGURE 38. Analysis of the range equation for the laser profiler.
INSTRUMENT SYSTEM 73

ture previously described (see "Laser Tracker" section). 

This refinement is termed a "last-pulse discriminator" 

and it is designed to cope with multiple-return signals 

(pulses) that will occur when flying the profiler over 

some types of vegetation; for example, trees or heavy 

brush. In a multiple-return sequence, the first pulse 

could represent a treetop; the next one or more pulses, 

individual branches; and the last pulse, the land surface. If the intent is to maximize the number of valid 

returns from the terrain, then a capability must be 

provided for identifying the critical last pulse.

In an over-simplified way, the last-pulse discriminator can be regarded as a split electronic circuit, one side 

of which features a pulse counter, and the other features a delay line created by 100 ft (30.5 m) of coaxial 

cable. The multiple-return signals are fed simultaneously through both sides of the circuit, and, thus, 

when the entire sequence emerges from the delay line, 

the counter already knows how many pulses there are. 

It is no great electronic feat to tell the signal processor 

which numbered pulse to act upon.

Foreseeable variations in environmental conditions 

during flight, as well as variations in data-collection 

objectives, prompted the deliberate inclusion of the following signal-processing options for range determinations with the laser profiler:

1. Last pulse, leading edge,

2. Last pulse, alternate between trailing edge and 

leading edge,

3. Last pulse, trailing edge, and

4. Last pulse, leading edge, alternate between first

pulse, leading edge.

The APT system operator selects the particular option 

desired by setting a rotary switch.

The original intent in designing the APT system was 

to mount the profiler on a gimbaled platform that could 

be slaved to the stable platform. In this way, the profiler could be held oblivious to the three-dimensional 

motion of the aircraft, so that its laser beam path would 

remain continuously oriented with the local vertical. 

When the need arose to invoke reasonable cost-saving 

measures, however, this feature was scrapped, and, in 

effect, the profiler now is hard mounted to the aircraft 

cabin floor. Within reasonable limits, corrections can 

and will be made for "off-vertical" laser ranges, inasmuch as the onboard computer will continuously store 

aircraft attitude data referenced to the stable platform.

Color Television Camera

When the APT system is used to collect terrain profile data, an important companion to the profiler subsystem is the color television (TV) camera subsystem 

(see fig. 25). Principal units in this subsystem are the 

camera, a video recorder, a control box for the recorder,

and a time-date generator. The original design for the 

first prototype APT system stipulated that the TV camera was to be bore sighted with the profiler and that 

both units were to be mounted on a gimbaled platform, 

which would be slaved to the stable platform of the 

IMU. Like the profiler, however, the TV camera fell 

victim to the cost-saving measures and is now hardmounted to the same rigid frame that supports the 

housing structure for the IMU and tracker.

The color TV camera chosen for the APT system is 

the Sony Model DXC-1640 (see fig. 39). The camera is 

housed in a metal case with suitable padding and handholds for operation and portability. Overall dimensions 

of the basic unit are about 9V4 in. (0.24 m) high, 7V2 in. 

(0.19 m) wide, and 17 in. (0.43 m) long, and it weighs 

nearly 10 Ib (4.5 kg). The power requirement is 11 W 

delivered as DC current at 12 V. Significant camera 

features include automatic exposure and gain controls 

for stable operation throughout a range in illumination 

("illuminance") from about 9 to 9,3000 ft-c (100- 

100,000 lux), normal operation throughout a range in 

environmental air temperature from 32°F to 104°F 

(0°C-40°C); National Television System Committee 

standard color reproduction through a 2/3-in. (26-mm) 

MF Trinicon tube pickup, and an individual picture 

(frame) comprised of a conventional raster of 525 lines, 

two-to-one interlace, and 30 frames per second, with 

300-line horizontal resolution.

The video recorder chosen for the APT system is the 

TEAC Video Corp. Model V-1000 AB-R (see fig. 40), 

which is a ruggedized cassette-load type. It can record 

simultaneously imagery from one camera and two 

audio sources, if desired. The unit is packaged in a 

metal box 6V4 in. (0.16 m) high, 9V2 in. (0.24 m) wide, 

and 13 in. (0.33 m) deep and is shock mounted in standard light-alloy racks. The unit weighs 23 Ib (10.5 kg), 

and the power requirement is 30 W delivered as unregulated DC current at 28 V. For normal operation, the 

recommended environmental temperature should be in 

the range from 5°F to 131°F (-15°C-55°C). Acutal powering, starting, and stopping of the recorder is through 

a separate control box, TEAC Video Corp. Model VS100G, which also contains certain appropriate test and 

check functions. The box is 4 in. (0.10 m) high, 97/8 in. 

(0.25 m) wide, and 6V2 in. (0.16 m) deep and weighs 

about 10 Ib (4V2 kg).

A standard video cassette, 73/i6 in. (0.18 m) wide, 

43/4 in. (0.12 m) deep, and 1V4 in. (0.03 m) thick accommodates magnetic tape l/2 in. (13 mm) wide and 

562 ft (171 m) long. Camera imagery is recorded on the 

tape by a two-headed device; its rotary motion, combined with the linear tape speed of 33/4 in./s (95 mm/s), 

results in a "helical" scan. Total recording time available per tape is 30 minutes; rewind time is about 2 

minutes. Because the recording heads should be re-
74 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

placed after about 500 hours of use, the recorder unit 

has an elapsed-time indicator that is direct reading to 

that limit.

The need to correlate camera imagery very precisely 

with profiler data means that time data, synchronized 

with the APT system master clock, must be superimposed on each image (frame). The time-date generator 

device selected to do this is a Laird Telemedia Inc. 

Model 3201 (see fig. 4LA). It is packaged in a metal case 

3 in. (0.07 m) high, 17 in. (0.43 m) wide, and 13V2 in. 

(0.34 m) deep and is mounted in standard light-alloy 

racks. The unit weighs about 10 Ib (4V2 kg), and its 

power requirement is 22 W delivered as single-phase 

current at 115 V and 400 Hz.

The electronic stratagem used to accomplish the desired time labeling of each camera frame involved modification of the time-date generator unit so that it could 

accept time-signal data from an external source, rather 

than rely on its own internal clock. These external 

time-signal data arrive as seven-decade binary-codeddecimal numbers, with a resolution of 0.01 second. The

camera imagery passes through the time-date generator so that each frame has the appropriate time-code 

number superimposed before the combination is fed to 

the tape recorder.

As the design of data-editing equipment evolved, it 

became apparent that it would be advantageous to 

have the abovementioned time-code labels on the audio 

track of the videotape. To accomplish this, an EECO 

Model MTG-551 time-code generator is connected in 

series with the Laird Telemedia unit. The EECO unit 

(see fig. 4LB) is packaged in a metal case !3/4 in. 

(0.04 m) high, 19 in. (0.48 m) wide, and 14 in. (0.36 m) 

deep and weighs 10 Ib (4V2 kg). Its power requirement 

is 10 W delivered as single-phase current at 115 V and 

60 Hz.

SPECIAL ROLE OF GRAVITY

Because the APT system, by virtue of its IMU, needs 

gravity information to perform its vital navigation 

task, it seems paradoxical to note that the APT system

FIGURE 39. Color television camera for the aerial profiling of terrain system.
SPECIAL ROLE OF GRAVITY 75

FIGURE 40. Video recorder and control box for the aerial profiling of terrain system.

also provides gravity information if certain conditions 

are satisfied. In view of this paradox and the importance of gravity information to studies of geophysics 

and geodesy, it seems appropriate to focus on the role of 

gravity in a single discussion here.

Perhaps the most fundamental principle of inertial 

technology is the fact that an accelerometer detects 

along the direction of its sensitive axis, two types of 

specific force or acceleration, each of which is indistinguishable from the other because they are manifestations of the same physical process: inertially referenced 

acceleration, associated with Newton's laws of motion, 

and gravitational acceleration, associated with mass 

attraction. The gravitational acceleration detected by a 

near-Earth IMU, such as the one in the APT system, is 

associated almost entirely with the mass of the Earth, 

although perceptible tidal effects are associated with 

the masses of the Moon and Sun. In terms of gravity, 

which is the sum of the Earth's gravitational acceleration and the centripetal acceleration linked to Earth 

rotation, such an IMU is said to detect a combination of 

Earth-referenced acceleration, whereby the inertial coordinate system is Earth centered and rotating at the 

Earth rate, and gravity. Thus, Earth-referenced acceleration can be obtained if gravity is known or estimated. Conversely, gravity can be obtained if Earthreferenced acceleration is known or estimated.

To extract Earth-referenced acceleration of the IMU 

for purposes of navigation, gravity can be estimated in 

two parts: a mathematically idealized normal part, associated with a reference ellipsoid having a size, shape, 

mass, and rotation rate similar to those of the Earth, 

and a disturbance part, associated with departures of 

the real Earth from the ellipsoid, especially those 

caused by irregular distributions of masses within the 

Earth. The normal part is represented by a geodetic 

mathematical formula; the disturbance part commonly 

is modeled using methods of probability and statistics.

To extract gravity sensed by the IMU, for purposes of 

geophysics and geodesy, Earth-referenced acceleration 

can be estimated using data from the laser tracker, 

which can serve as an independent navigator not affected by gravity. The tracker provides a time history of 

aircraft position while a retroreflector is in view and 

thus provides velocity and acceleration data, given sufficient viewing time or a sufficient number of successive passes within viewing range of the retroreflector. 

By estimating and removing a linear combination of 

error sources tied to the IMU and tracker, aircraft motion can be effectively eliminated. Thereby, the redundant navigation permits extraction of the gravitydisturbance vector, the magnitude and direction of 

which correspond closely to gravity anomaly and deflection of the vertical.
76 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

Gravity also plays a special role in error analysis of 

the APT system. In the design engineering phase of the 

APT system development, it was recognized that errors 

associated with accelerometer and gyro drift cannot be 

distinguished from horizontal rates of change of gravity magnitude and deflection of the vertical. Further 

analysis indicated that, because accelerometer and 

gyro drift are time correlated in contrast to the gravity

A TIME DATE GENERATOR

B TIME CODE GENERATOR

FIGURE 41. Time-date and time-code generators for the aerial 

profiling of terrain system.

effects which are space correlated, it would be possible 

to separate the inertial and gravity effects by flying the 

same path in opposite directions during short time periods. Such a procedure would permit the APT system to 

develop a gravity gradient map to the extent that the 

retroreflectors serve as points between which straight 

lines in gravity gradients can be drawn.

From a broader perspective, it is recognized that 

gravity irregularities induce acceleration errors in the 

IMU which are a major source of position error. These 

errors are defined as the difference between the true 

gravity vector at the aircraft and the gravity vector 

which is computed on board using the postulated gravity model as a function of the IMU-indicated aircraft 

position. The total gravity-induced error can be expressed as a sum of two terms: Schuler feedback, which 

gives rise to oscillatory error change in the horizontal 

channels of the accelerometer triad and exponential 

error growth in the vertical channel of the triad, and 

the error in modeling true gravity. The second term, 

which is a challenging subject of ongoing research, can 

be considered the sum of an acceleration bias, a linear 

combination of gravity-disturbance states, and a noisy 

residual. The success of gravity modeling depends 

largely upon how accurately gravity disturbance can 

be estimated by applying principles of probability and 

statistics to the term involving gravity-disturbance 

states.

Finally, it should be noted that it is possible to separate the effects of gravitation and inertia on the basis 

of the structure of their respective fields. Such a separation is possible for the second and higher derivatives 

of the gravity potential but not for the gravity vector 

itself. In practice, the separation may be made by 

simultaneously measuring the first and second derivatives of the gravity potential, using an accelerometer in 

combination with an airborne gravity gradiometer. For 

an aircraft with an inertial stable platform, the second 

derivatives of the force field do not contain inertial 

disturbances, and this enables measurement of purely 

gravitational second-order gradients. The gradients 

are integrated along the flight path to yield the gravitational force vector. It is of interest that the first airborne gravity gradiometer is currently (1985) under 

development, which offers hope that deployment of the 

APT system with such a gradiometer may soon be feasible.

AIRCRAFT FOR DEPLOYMENT OF 

INSTRUMENT SYSTEM

Among the operational specifications is the stipulation that the APT system be deployable in a relatively 

light fixed- or rotary-wing aircraft. "Relatively light" is
AIRCRAFT FOR DEPLOYMENT OF INSTRUMENT SYSTEM 77

intended to mean below the Federal Aviation Administration's (FAA) gross-weight classification limit of 

12,500 Ib (5,670 kg). Because deployment in a fixedwing aircraft presents the greater design and operating 

challenge and because most field applications envisioned for the APT system will favor using that type 

aircraft, the following discussion ignores rotary-wing 

aircraft. Nothing in the design effort has precluded in 

any way, however, deployment in rotary-wing aircraft 

should the field situation so dictate.

To help in identifying the most suitable aircraft 

makes and models, the list of selection factors shown in 

table 3 was developed. Some data in the table that 

concern the APT system, crew, and aircraft modifications are shown for the originally intended system design and for the modified prototype design that will be 

flown first. Comparison of these particular data reveals 

the extent to which some of the original design objectives were necessarily compromised to accommodate 

budget constraints.

TABLE 3. Factors bearing upon aircraft selection 

Flight mission factors

Arbitrary ground speed 120 mi/hr or 104 knots (193 km/

hr) = 175 ft/s (52 m/s).

Maximum round trip flying time to field-survey site 2 hours.

Calibration and terrain-profiling time 2-1/4 hours.

Loiter time and reserve 3/4 hour.

Maximum total flight time 5 hours.

Standard rate turns during profiling time 37s.

Normal operating flight altitudes above terrain less than 3,000 ft

(914 m).

APT system and crew factors

APT system size 15.5 ft3 (0.43 m3) ----------- 64 ft3 (1.81 m3)*.

APT system weight 400 Ib (182 kg) --------1,470 Ib (667 kg)*.

APT system power load 1,000 W--------------------4,800 W*.

Maximum crew of four persons----------------- 680 Ib (309 kg).

Aircraft modification factors

Window in bottom of fuselage for tracker, profiler, and TV camera 

operation, 33 in. (0.84 m) x 53 in. (1.35 m)-------65 Ib (29 kg)

Base structure, secured to cabin floor beams to mount APT system 

components -----------------------------------45 Ib (20 kg)

Installation of two alternators to meet electrical power load 100- 

amp capacity each.

"Installation of gas-turbine auxiliary power unit, inverters, battery, 

and cabling----------------------------------690 Ib (313 kg)

"Installation of cabin air-conditioning system -------75 Ib (34 kg)

Aircraft performance factors

Good short-field takeoff and landing characteristics.

Good slow-flight characteristics.

Good margin on payload capability.

Economical operating and maintenance characteristics.

No cabin pressurization or air conditioning.

*Modified design, APT system.

The aircraft review task was narrowed quickly by 

the particular mission factors of relatively low flight 

speeds and altitudes. These factors, coupled with the 

desired low initial, operating, and maintenance costs, 

immediately favored piston-engine type aircraft rather 

than gas-turbine or turboprop types. The latter conclusion holds, however, only if the intent is to field an APT 

system built in accordance with the original design. An 

added selection factor concerned the requirement for 

openings through the cabin floor. This inclined the review task toward high-wing-type aircraft to avoid the 

need for complicated modifications in wing-root structure.

With the range of possible aircraft choices effectively 

limited through application of the foregoing choice factors, Survey and Draper Laboratory personnel now 

could introduce such secondary factors as cabin size 

and available FAA-approved designs for cabin-floor 

openings. In due course, this round of evaluation led to 

final choices of the following twin-engine aircraft:

1. North American Rockwell Commander, Model 

680FL, gross weight 8,500 Ib (3,860 kg), with two 

Lycoming Model IGSO-540-BIA piston engines, 

six cylinder, supercharged, fuel injected, 360 

horsepower (hp) each (maximum continuous) at 

3,200 rpm. (Suitable for fielding the APT system 

in its "original design.")

2. DeHavilland Twin Otter 300, Model DHC-6, gross 

weight 12,500 Ib (5,670 kg), with two Pratt and 

Whitney Model PT6A-27 gas-turbine engines, 

620 shaft hp each. (Suitable for fielding the APT 

system in its reduced-cost or "modified" design.) 

Inasmuch as the first prototype model of the APT system to be flown has been built to the modified design, 

the remaining discussions in this paper concern only 

the DeHavilland Twin Otter aircraft.

The principal performance characteristics of the 

Twin Otter aircraft are assembled in table 4, and its 

overall appearance is as shown in figure 42. The general arrangement of APT system components, as 

mounted in the aircraft cabin, is in the artist's cutaway 

perspective (fig. 43).

OPERATION OF THE INSTRUMENT SYSTEM

During a flight mission, the APT system can be used 

for scientific measurement in three principal ways that 

relate to the profiling of terrain, establishment of horizontal and vertical control, and control or monitoring of 

the position and orientation of a companion remote 

sensing instrument. Regardless of the way in which the 

APT system is to be used, its operation requires that 

three retroreflectors, centered over three presurveyed 

ground-control points, be tracked individually during
78 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

TABLE 4. Performance characteristics, DeHavilland Twin Otter aircraft

POWER PLANTS 

Pratt & Whitney PT6A-27, 620 shaft horsepower each.

PERFORMANCE CONDITIONS

Performance estimations are based upon U.S. Standard (1962) atmospheric conditions and performance is contingent upon engine manufacturer's guaranteed performance as indicated in FAA Type Certificate.

OPERATING SPEEDS AT 12,5000 POUNDS

Maximum speed ----------------------------

Cruise [10,000 ft (3,050 m)], 70-percent power 

Stall (clean configuration) ------------------

Stall (landing configuration) ----------------

Minimum single-engine control speed --------

Miles per hour Knots Kilometers per hour

207 182 333

170 150 274

84 74 135

66 58 106

68 60 109

MAXIMUM CRUISING RANGE

(Includes fuel for taxi, takeoff, climb, normal cruise, descent, and landing, and 45-minute reserve at optimum altitude.) 

A. With maximum fuel full tanks (2,470 Ib) ....................................................-. -1680 statute mi (1,090 km)

(1,120 kg) (590 nmi)

B. With special equipment and operating crew, and fuel available up to gross weight *450 statute mi (720 km) 

(390 nmi)

Twin-engine initial rate of climb (10° flaps) ------------------------------------------------------------ 1,600 ft/min (490 m/min)

Single-engine initial rate of climb (0° flaps) ------------------------------------------------------------ 340 ft/min (100 m/min)

Twin-engine service ceiling --------------------------------------------------------------------------- 26,700 ft (8,140 m)

Single-engine service ceiling -------------------------------------------------------------------------- 11,600 ft (3,540 m)

Takeoff distance to clear 50-ft (15-m) obstacle ---------------------------------------------------------- 1,500 ft (460 m)

Maximum gross weight 12,500 Ib (5,670 kg).

Empty (operating) weight to be determined after installation of special surveying equipment. 

Approximate only.

the flight mission. Additional retroreflector-equipped 

ground-control points may be needed in the project area 

to satisfy the requirement that a target be visible for 

tracking at 3-minute intervals during the survey 

flight. In this event, the added unsurveyed retroreflectors also are tracked individually during what might be 

called an initialization flight, before the collection of 

the intended survey data. Approximate coordinates are 

needed for the unsurveyed retroreflectors to ensure 

successful tracker search. They are obtained by measurements on such large-scale maps as the Survey 

standard 7 1 /2-minute quadrangle sheets (scale 

1:24,000).

For the terrain-profiling use, the APT system is 

flown along a course that overlies those lines on the 

ground for which profile data are desired. As it traverses that course, the APT system maintains a threecoordinate scheme of reference (latitude, longitude, 

and elevation), with which the laser profiler range 

measurements to the land surface are combined vectorially. The combination is vectorial in the sense that 

aircraft roll and pitch adjustments are made to convert 

range measurements to the desired vertical direction. 

This produces a profile of the terrain along a line that 

ideally lies vertically beneath the aircraft flight path.

Moderate roll motion of the aircraft, however, will 

cause moderate departures from this ideal line, inasmuch as the profiler is hard mounted to the cabin floor. 

The video subsystem, which operates concurrently 

with the profiler, scans and records an aerial view (in 

color) of the terrain path that is being profiled. Such 

video imagery serves an important postmission role in 

the editing of the profile data to remove unwanted laser 

signal returns from foliage and manmade structures 

and to verify that the data have indeed been taken 

along the correct path. It also serves effectively in the 

precise location or placement of the terrain profiles on 

aerial photographs or large-scale maps.

For the establishment of horizontal and vertical control, the APT system determines the locations of desired unsurveyed ground-control points, as marked by 

retroreflectors. Normally, in this type of field use, the 

laser profiler and video subsystem are not operated. 

The fight technique simply entails an attempt to follow 

the preplanned sequence of individually tracking all 

retroreflectors mounted on the presurveyed and the unsurveyed ground-control points. The tracking sequence 

might conceivably feature a crisscrossing of the entire 

project area. In the postmission processing of the flight 

data, adjustments to the inertial navigator data are
OPERATION OF THE INSTRUMENT SYSTEM 79

founded on tracker measurements to the retroreflectors 

at the presurveyed ground-control points. From this 

foundation, the further processing and adjusting of the 

data yield the position coordinates for the unsurveyed 

ground-control points. High-accuracy results can be expected, inasmuch as the unsurveyed sites are exclusively for tracked retroreflector targets.

For the control or monitoring of the position and orientation of a companion remote sensing instrument, 

the APT system normally is operated without the laser 

profiler and video subsystem. As with all other uses, it 

maintains a three-coordinate scheme of reference, as it 

controls or monitors the orientation of the companion 

instrument with reference to the stable platform in the 

inertial navigator. The companion instrument can be 

any one of a number of devices, such as an aerial camera, an infrared scanner, magnetometer, or sidelooking radar. The usefulness of the remotely sensed 

data collected by such instruments is greatly enhanced 

by continuously knowing the instrument position and, 

for certain instruments, the orientation of their receiving optics.

GROUND CONTROL

Field operation of the APT system presupposes some 

beginning knowledge of where things are in the project 

area; that is, the latitude, longitude, and elevation coordinates for at least three ground-control points. 

These coordinates are to be determined and given in 

terms of the national control networks; for the United 

States, these are the NAD of 1927 (for latitude and 

longitude) and the NGVD of 1929 (for elevation). The 

obvious intent is to ensure that data collected by the 

APT system are compatible with existing maps and 

other geographic information systems. The groundcontrol points also serve to orient and scale the flight 

survey. These attributes emerge from the postmission 

data processing and adjustment when, in the final 

steps, the survey is fitted to the ground control. As a 

sound and practical operating technique, the combined 

array of presurveyed and unsurveyed ground-control 

points should be designed to embrace the entire survey 

area so that the APT system always is used as an interpolative rather than extrapolative positiondetermining device.

FIGURE 42. DeHavilland Twin Otter aircraft.
80 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

The national control networks consist of bench 

marks and triangulation stations; the former are 

points whose elevations are accurately known, and the 

latter are points whose latitudes and longitudes are 

accurately known. Over 1 million bench marks dot the 

48 contiguous States, and most are along highways and 

railroads at intervals of about 2 mi. On standard Survey quadrangle maps, a bench mark is shown as a 

small black cross, and just a brief inspection of a given 

map shows that most are found in the valleys and 

places where the roads are concentrated. Triangulation 

stations are not as plentiful and number only about 

200,000. They are shown as black triangles on standard 

Survey maps and generally appear on high ground, 

such as hill and mountain tops.

Typically, a field project that is to be surveyed by the 

APT system will be in relatively flat country or along 

a stream. Thus, because the odds favor the preexistence 

of a number of bench marks in the project area, the 

desired elevation control is rather easily established. 

Triangulation stations in the project area are likely to

LINE PRINTER 

COMPUTER MODULES

MAGNETIC TAPE RECORDER 

INVERTERS AND SYSTEM V|DEQ RECORDER

VIDEO CONTROL BOX 

MAGNETIC TAPE RECORDER

be scarce to nonexistent, however, and a field-survey 

party must then establish a minimum of three that 

appropriately bracket the area.

The amount and type of local vegetation and topographic relief will strongly influence the type of surveying instruments used to establish triangulation stations. If the intervisibility among the proposed station 

locations is good, then the least costly surveying technique is the traverse method whereby plane angles are 

measured with theodolites and distances with 

electronic-distance-measuring equipment. If station 

intervisibility is poor, then semiautomatic satellitetracking instruments will be used; for example, the 

JMR-1, the Magnavox 1502, or the Motorola MiniRanger. The ideal technique in that event is to operate 

several instruments simultaneously at several stations, one of which must be the nearest preexisting 

triangulation station. In this manner, appropriate 

"relative" position data are measured, and the necessary transformations can be made to the desired NAD. 

A day of satellite tracking will yield horizontal position

PLASMA DISPLAY 

TERMINAL

OPERATOR'S 

KEYBOARD,

CABIN AIR CONDITIONER AND AUXILIARY 

POWER UNIT IN AFT BAGGAGE COMPARTMENT

REMOTE CONTROL PANEL 

MAGNETIC TAPE CONTROLLER 

COMPUTER CONTROL PANEL 

DC POWER SUPPLY 

TIME-CODE GENERATOR 

ELECTRONICS UNIT

k TIME-INTERVAL METER 

' REFRIGERATOR (IMU COOLANT) 

k FAN CONTROLLER (IMU) 

^DC POWER SUPPLY

k LASER SIGNAL PROCESSOR AND MASTER CLOCK 

'PROFILER 

'VIEWING WINDOW

FIGURE 43. Arrangement of aerial profiling of terrain system components in DeHavilland Twin Otter aircraft.
OPERATION OF THE INSTRUMENT SYSTEM 81

coordinates accurate to within 1 m. When the NAVSTAR satellite navigation system, now being developed by the Defense Department, becomes operational, 

even greater accuracy will be possible with a tracking 

time of only 2 hours. To be compatible with most uses 

for the APT system, the field surveys for triangulation 

stations should be held to third-order standards. However, when the APT system itself is to be used to establish horizontal geodetic control, the initial field surveys 

should be held to second-order standards.

Despite the general profusion of bench marks, they 

will seldom be found at the exact sites chosen for 

ground-control points. Thus, short lines of levels must 

be run to transfer elevations from the nearest bench 

marks to all the triangulation stations established as 

ground-control points and to as many intermediate 

control points (when used) as practical. Normally, such 

lines will be run to third-order standards. For especially precise uses of the APT system, however, level 

instruments and procedures appropriate for secondorder standards should be used.

FLIGHT PLANNING

Thorough advance planning for field work is instinctive in the Geological Survey. Beyond the innate incentives, however, the use of the APT system has an added 

incentive because its operating cost (aircraft, instrument maintenance, crew salaries and per diem, data 

editing, and computer processing) will be high, about 

$5,000 per flight hour. Success of a flight mission, as 

well as its most efficient execution, will hinge on thorough and astute flight planning.

A well-designed plan for a survey mission with the 

APT system literally means the detailed layout of the 

entire flight path from the moment of departure from 

the principal airport nearest the field-survey area until 

the moment of return. That path must include, therefore, the route to be navigated to the survey area, the 

precise pattern to be flown initially over all retroreflectors mounted on ground-control points in the area, the 

precise pattern to be flown to collect the desired aerial 

survey data, and the route to be navigated back to the 

airport of origin. Layout of the flight path on a suitably 

scaled map will graphically illustrate the sequence of 

flight maneuvers needed to accomplish the mission. 

From the map layout, the horizontal and vertical coordinates of each retroreflector and the horizontal coordinates for the beginning and ending points of each discrete segment of the flight path (for example, each 

terrain profile) will be scaled and listed. That list also 

will include the scaled position coordinates for a convenient point on the aircraft parking ramp at the departure airport. Advance study of these listed data and the 

map layout by the flight crew must be sufficient to

ensure mission performance without hesitation and 

preliminary familiarization flights.

Design of the most suitable flight path is dominated 

by the following criteria:

1. At 3-minute intervals during the aerial survey part 

of the mission, a retroreflector must be tracked by 

the laser tracker to update the APT system consistent with a 1/2-ft vertical accuracy. For a flight 

speed of about 120 mi/hr or 104 knots (193 km/hr), 

this suggests a retroreflector spacing of about 6 mi 

(10 km) along the flight path. If vertical accuracy 

is relaxed, then fewer updates are needed and 

retroreflector spacing can be increased; for example, 1-ft (0.30-m) vertical accuracy requires updates at about 4-minute intervals, or a retroreflector spacing of about 8 mi (13 km); 2-ft (0.61-m) 

accuracy equates to 6-minute updates and a spacing of about 12 mi (19 km).

2. Aircraft turning maneuvers should not exceed socalled standard-rate turns, in which the limiting 

bank (roll) is 15°. At that limit, the turning rate 

(change in azimuth) is 3°/s, or 2 minutes for a 

complete circle. This protects the stable member 

of the IMU from undue centrifugal acceleration. 

Flight experience ultimately may allow some relaxation in this criterion.

The salient features of a flight plan are shown in 

figure 44, for an aerial survey related to a flood-plain 

mapping problem. The illustration covers what might 

be termed one module of the flight plan, inasmuch as it 

spans only a 3.5-mi (5.6-km) portion of the much longer 

reach of the Framington River that is to be surveyed. 

This flight-plan module typifies, however, the kinds of 

maneuvers involved in collecting terrain-profile data 

along a stream valley. The modular layout can be progressively repeated, with minor local variations, as 

many times as needed until the entire stipulated reach 

of the river is covered.

The flight-path layout in figure 44 is based on a constant aircraft speed of 115 mi/hr or 100 knots (185 km/ 

hr), a flight altitude of 2,000 ft (610 m) above the terrain, and all turns made at standard rate. Each loop in 

the path is sized so that it represents a flying time of 

about 3 minutes. This can be verified by following the 

consecutively numbered path segments. For convenience in layout, a segment length unit was arbitrarily 

chosen as the distance traveled in a 45-second flight. 

This means that four consecutive units represent a 3- 

minute interval. As the illustration shows, the retroreflectors are logically positioned near the convergence of 

flight paths to insure that tracking can be done at the 

required 3-minute intervals throughout the flight. 

However, at this flight altitude, the tracker can only 

see a retroreflector when the aircraft is within a horizontal range of about 3,200 ft (975 m). A circle of this
82 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

RETROREFLECTOR AT GROUNDCONTROL POINT

NUMBERED 45-SECOND 

INTERVAL ALONG FLIGHT PATH. 

HEAVY DASHES DENOTE WHERE 

TRACKING OF NEAREST 

RETROREFLECTOR OCCURS

400'/ '/ / /

/ //& .!/

^ >-1i»! *r±~^ i if

.© ( i / A

FIGURE 44. Flight path layout for profiling the Farmington River flood plain near Avon, Conn.
OPERATION OF THE INSTRUMENT SYSTEM 83

radius is shown (fig. 44) around each retroreflector, and 

the flight-path segments that fall within any circle represent tracking opportunities.

The location and orientation of the desired streamvalley profiles are shown in figure 44, and the order in 

which they actually will be flown is given by the sequence of circled lower-cased letters a through k.

FLIGHT MISSION

Presentation of an abbreviated script for a hypothetical flight mission with the APT system is a convenient 

way to highlight the principal operating techniques. 

For this mission, it is assumed that the aerial survey is 

to collect terrain-profile data along specified paths in 

the broad flood plain of a fairly sizable stream valley. 

The sequence of operating steps is as follows:

1. On a convenient date before the flight, an advance 

ground party mounts a retroreflector on each 

ground-control point that is to be used by the APT 

system. The party also reaffirms the availability 

of appropriate ground-power equipment at the selected principal airport nearest the field-survey 

area. This equipment must support the APT system in all the preflight and postflight ground 

phases of the mission.

2. On the day before the flight mission, the APT system is flown to the selected airport near the fieldsurvey area. The aircraft is taxied to and parked 

over the preselected point on the ramp, whose position coordinates already have been scaled from a 

map and listed. The APT system is connected to 

the ground-power equipment, and, at the appropriate number of hours before the flight is due to 

begin, the operator sequences the system through 

STARTUP AND TEST and into CALIBRATE 

AND ALINE. While the system is in this latter 

mode, the operator loads his list of position coordinates for all retroreflectors and the start and end 

points for each discrete flight-path segment into 

the computer. He also verifies that suitable 

gravity-model and reference-ellipsoid data are 

available for the survey area, which will be used 

in postmission data processing. A few minutes before flight time, the APT system is switched to 

STANDBY, the aircraft auxiliary power unit 

(APU) is started, and the ground-power equipment is disconnected.

3. Just before the aircraft is ready to taxi as the flight 

time nears, the APT system is switched to EN 

ROUTE. The inertial navigator now will measure 

and record the actual path of the aircraft as it 

moves into and through the beginning phases of

the flight mission. Approximately 30 minutes 

may elapse in this mode before the APT system 

arrives within range of the first retroreflector in 

the survey area.

4. As the aircraft nears a point where the tracker will 

be within range of the first listed presurveyed 

retroreflector, the APT system is switched to 

SEARCH. When the tracker has found and locked 

onto the retroreflector, the system is switched automatically to AERIAL CALIBRATE. The preplanned calibration segments of the flight path 

are now flown so that each retroreflector in its 

proper turn can be tracked individually. The 

usual pattern is to track each presurveyed 

retroreflector in one continuous pass, followed by 

a return pass in which each is tracked again. At 

some convenient time during this round-trip calibration exercise, each unsurveyed retroreflector 

is tracked, and its measured and computed 

position-coordinate data stored for subsequent use 

in the aerial surveying part of the flight mission 

as well as in postmission data processing.

5. When the calibration exercise is complete, the APT 

system is switched to AERIAL SURVEY, and the 

preplanned segments of the flight path needed to 

collect the specified terrain profiles are flown. 

During this part of the flight mission, individual 

retroreflectors of opportunity are tracked so that 

gaps in tracker data do not exceed 3 minutes. The 

tracking procedure is automatic and in keeping 

with the flight plan. All tracker data are recorded, 

together with the terrain-profile data, for use in 

postmission data processing to update the 

inertial-navigator position and velocity information. The only such updating during the flight 

mission itself occurs whenever the position error 

is determined to have grown to more than 200 ft 

(61 m). The time and magnitude of each inflight 

update are also recorded for use in postmission 

data processing.

6. When the desired terrain-profile data have been collected, the APT system is switched once more to 

AERIAL CALIBRATE, and a final round-trip calibration exercise is flown. This will allow the subsequent (postmission) prorating of any changes in 

APT system performance parameters that may 

have occurred during the flight mission. When 

the calibration exercise is complete, the APT system is switched to EN ROUTE, and the aircraft 

returns to its home-base airport. 

The products from the flight mission are two sets of 

magnetic tapes, one of which holds the tracker and 

terrain-profile data and the other, the color video imagery of the terrain-profile paths.
84 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

MAINTENANCE AND SERVICE

The inertial instruments in the APT system must be 

powered continuously to achieve their peak highaccuracy performance. Whenever a shutdown occurs, 

several days of warmup and retesting are necessary to 

restabilize the internal temperature gradients at the 

proper levels and to remeasure the performance 

parameters of all inertial sensors. Unexpected breakdowns will cause substantial delays, interrupt carefully drawn flight schedules, and, thus, be intolerably 

costly. For these reasons, periodic preventive maintenance shutdowns are deliberately built into the overall 

field scheduling for the APT system.

During a maintenance shutdown, the IMU and all 

other instruments will be opened and inspected; components that have been performing marginally will be 

replaced. Routinely, the poorest performing gyro and 

accelerometer will be replaced with better performing 

units from the supply of spares. The units that are 

pulled from the IMU will be reconditioned, recalibrated, and returned to the spare supply. The ideal 

frequency of such maintenance shutdowns can be determined only from experience. At the outset, once 

every few months will be tried.

Records will be maintained of APT system performance parameters and operating temperatures as measured during normal field operations. Sudden changes 

in any parameter will indicate a potential for an imminent problem that should be promptly investigated and 

overcome. Gradual changes in any parameter will 

point toward a longer range problem and might, therefore, serve to gage the time for a maintenance shutdown.

POSTFLIGHT DATA PROCESSING

The basic purpose of data processing after the flight 

mission is to reconstruct as accurately as possible the 

three-coordinate position information for the aircraft 

at specified time intervals along the flight path. If profiler data also were collected, then these can be combined with the processed position information to produce terrain profiles.

The kinds of "raw" data brought back (on magnetic 

tape) from a terrain-profiling flight mission include 

color video imagery of the profile paths, profiler ranges, 

tracker ranges, tracker gimbal angles, IMU gimbal angles, and accelerometer readouts. For a 3-hour mission, 

this could total as much as 10 megabytes of data, to 

which there would be added about 1 megabyte of lowerfrequency data comprised of system initialization, calibration, and related information.

A schematic of how these flight data would flow 

through the postmission data-processing software is 

shown in figure 45. Principal steps in the process are

numbered in the appropriate boxes in the figure and 

are briefly described as follows:

1. Data editing. A program that transfers magnetic tape contents (the flight data) onto a disc in a way 

that reformats from DEC-compatible to IBMcompatible format. The program sorts, prints, and plots 

raw data; removes or replaces "outriders" in the data; 

and reformats as appropriate for the next steps.

2. Renavigating. A program that solves the appropriate navigation equations, using the recorded system 

initialization, calibration, and accelerometer data to 

reconstruct, at 80-ms intervals, inertial-navigatorindicated values of aircraft velocity (relative to the rotating Earth) and position coordinates. Concurrently, 

the program computes tracker-indicated values of aircraft velocity and position coordinates using the 

recorded measurements of tracker range and IMU and 

tracker gimbal angles. These computed values are subtracted from the corresponding inertial-navigatorderived values to obtain "difference" data.

3. Data compressing. A program that accepts the 

renavigation difference data (step 2) as input and performs a polynomial curve fit on each successive group 

of 25 values, which is equivalent to a 2-second time

REALTIME 

POSITION AND 

PROFILER DATA

(AFTER HURSH, 1984.)

FIGURE 45. Information flow through postmission dataprocessing software.
PERFORMANCE EVALUATION FLIGHTS 85

interval. At each 2-second time boundary on the fitted 

curve, the polynomial is evaluated for output to the 

Kalman filter. The program also selects the corresponding values of inertial-navigator-indicated aircraft acceleration, velocity, and position coordinates, 

tracker-indicated range and position coordinates, and 

IMU and tracker gimbal angles for output to the 

Kalman filter. The net results of compressing the data 

are to reduce its volume by a factor of 25, average out 

errors, and synchronize, in 2-second time steps, all 

parameter values that are to be fed to the Kalman 

filter.

4. Kalman filtering. A program that uses an optimal estimation algorithm (mathematical procedure) to 

produce a 2-second time sequence of best estimates of 

the state vector of the APT system (see "Software" section, Calibrate and Aline mode, for a general description of Kalman filtering). Inputs to the Kalman filter 

include data from step 3 and available gravity data. A 

detailed and rigorous filter description is beyond the 

scope of this paper. The interested reader may consult 

Soltz and others (1981, p. 12 15) for a mathematical 

description. Suffice it to say that the filter output embodies error estimates that propagate forward in time 

for principal elements of the APT system state vector 

which include: 

Elements Description

3 Errors in indicated IMU position coordinates

3 Errors in indicated IMU velocity

3 Stable platform misalinement angles

3 Accelerometer bias errors

3 Gyro drift-rate bias errors

4 Gimbal angle readout bias errors 

1 Tracker range bias error 

3 Survey errors at each retroreflector 

3 Gravity disturbance errors at each 

retroreflector

5. Back filtering. A program that uses an optimal 

estimation algorithm similar to that used in the 

Kalman filter. The program processes the same data 

(as in step 4) in reverse time sequence to produce error 

estimates that propagate backward in time.

6. Optimal smoothing. A program that optimally 

combines the two sets of time-sequenced error estimates from steps 4 and 5. This program produces a time 

history of the best possible estimates of all errors in the 

system at 2-second intervals.

7. Recombining. A program that accepts the time 

history of errors from step 6, interpolates the error values to yield a 40-ms time sequence, and combines these 

results with the appropriate inertial-navigatorindicated position data from the renavigation program 

(step 2) to produce a 40-ms time sequence of best estimates for the aircraft position coordinates. For stipulated time periods, these results now may be combined 

with the profiler range data, appropriately adjusted for

pitch and roll of the aircraft, to yield the desired survey 

data; that is, the terrain profiles.

PERFORMANCE EVALUATION FLIGHTS

OVERALL OBJECTIVES

The logical and necessary conclusion for the long gestation period for the prototype APT system was the 

successful accomplishment of certain specifically defined proof-of-concept, or performance-evaluation, 

flights. The Survey drew these specifications to ensure 

the collection of flight data that would permit a quantitative measure of the following:

1. Terrain-profile accuracy under selected spacings 

and geometries of retroreflector arrays.

2. Terrain-profile accuracy under selected aircraft maneuvers, such as standard-rate versus halfstandard-rate turns.

3. Terrain-profile accuracy for a variety of topographic 

features that includes ponds, streams, densely 

wooded tracts, urban areas, and farmland.

4. Accuracy in determining the positions of unsurveyed retroreflectors under selected spacings for 

the initial presurveyed or fixed control points.

CALIBRATION RANGE

To meet the stated objectives for the performanceevaluation flights, a large amount of precisely measured ground truth was needed. No land area featuring 

such ground truth existed within easy flight range of 

the aircraft operations base at Hanscom Field, Bedford, 

Mass. Thus, the Survey chose to design and prepare its 

own "calibration range" (see fig. 46) in an area roughly 

10 mi (16 km) wide (east and west) and 30 mi (48 km) 

long (north and south). Hanscom Field lies virtually at 

the midpoint of the eastern boundary, and the range 

area embraces the Massachusetts' cities of Framingham and Lowell, near the southern and northern 

boundaries, respectively. Western fingers of the 

greater Boston metropolitan area reach a few miles 

into the eastern part of the area.

The calibration range design was based on standard 

Survey 71/2-minute quadrangle maps (outlined and labeled in fig. 46), quadrangle-centered aerial photography, and geodetic control diagrams. The sites for 15 

control points or stations were selected in locations 

scattered throughout the range area from Lowell to 

Framingham (fig. 46). Precise field surveys established 

the three-coordinate position data for each point, and a 

retroreflector was mounted over each before the evaluation flights. The point named "Haystack" (fig. 46), 

which was established some years ago by the National 

Geodetic Survey, was included in the survey net and 

provided the corrections to the NAD of 1927.
86 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

A Doppler satellite relative positioning technique 

was used to establish field positions. Four Magnavox 

Model 1502 receivers were set up at the first four control points to be surveyed, which were chosen to form a 

quadrilateral. The receivers simultaneously observed 

the required number of satellite passes over a period of 

2 to 3 days. When sufficient data were recorded, two 

receivers were moved to two new points, which were 

chosen to form a new quadrilateral, and a new data set 

collected. This receiver leap-frogging process was repeated until data for a series of interlocking and over71 °30' 

42"45'

lapping quadrilaterals covered the entire area. From 

the lines drawn to connect the points shown in figure 

46, the general scheme of quadrilaterals can be visualized. Because of radio interference, the point named 

"Lab" was not occupied with a Magnavox receiver. The 

point named "Lincoln" was occupied as an alternative, 

and, once its position coordinates had been determined, 

they were transferred to the Lab site by standard surveying techniques.

Second-order leveling was used to determine elevations at all control points, relative to the NGVD of

EXPLANATION

CITY

CONTROL POINT

WITH RETROREFLECTOR

SURVEY CONTROL POINT

HAYSTACK

LITTLETON

42°30'

MARLBOROUGH 

*

FRAMINGHAM 

FRAMINGHAM

10 MILES 

I

42 ° 15 ASHLAND 

FIGURE 46. Control diagram and retroreflector sites for calibration range.
PERFORMANCE EVALUATION FLIGHTS 87

1929. Elevations also were determined for the local 

control needed for large-scale aerial photography flown 

at six of the retroreflector sites Acton, Ashland, 

Framingham, Littleton, Nagog, and Sudbury. From 

this photography, large-scale (1:800) photomaps with a 

1-ft contour interval were prepared for use in evaluating the accuracy of profiles measured by the APT system. Individual map coverage is a square approximately 1,570 ft (480 m) on a side.

Each of the 15 retroreflector sites is described in detail in station descriptions compiled by the field survey 

crew. Ultimate life and recoverability are further protected by a bronze bench mark tablet, installed in a 

permanent location at each site.

Gravity measurements were made throughout the 

calibration range area at 161 stations, that included 14 

of the retroreflector sites. The field party used a La 

Coste and Romberg geodetic gravity meter, and the 

resultant 161 site-measurement values are to be combined with 185 values extracted from gravity files 

maintained by the National Oceanographic and 

Atmospheric Adminstration in Boulder, Colo., to compile a gravity map for the range area.

Results of the foregoing survey and gravitymeasurement work are summarized in table 5 for the 

15 retroreflector sites.

SPECIFIED FLIGHTS

Four distinct flight plans were designed to meet the 

specified performance-evaluation objectives. The 

flights were to be made at an altitude of 2,000 ft (610 m) 

above mean ground level and at a constant speed of 115 

mi/hr or 100 knots (185 km/hr). The intent was to fly

directly over the retroreflectors. Details of each flight 

are summarized as follows: 

Flight 1

Purpose: To determine accuracy of the APT system 

for point positioning surveys.

Flight plan: Layout as shown in figure 47. Retroreflectors are to be tracked in the indicated 

sequences, after departing Hanscom Field. The 

APT system is to be initialized by tracking, in sequence, the Mohawk, Acton, Littleton, and 

Haystack retroreflectors. Thereafter, a southbound flight is to be flown with a northbound return first over the retroreflectors connected (fig. 

47) by the solid line (loop 1) and then over those 

connected by the dashed line (loop 2). The return to 

Hanscom Field is over the same retroreflectors (in 

reverse order) used to initialize the APT system.

Aircraft turns: All are to be made at one-half the

standard rate; that is, lV2°/s. 

Flight 2

Purpose: To repeat flight 1 on a different day, after 

a new preflight calibration and alinement procedure for the APT system. This should illustrate the 

accuracy with which the APT system can repeat 

surveys or detect changes in control-station position (subsidence). 

Flight 3

Purpose: To determine accuracy of the APT system 

in terrain profiling, with tracker updates at about 

5-minute intervals.

Flight plan: Layout as shown in figure 48. Retroreflectors are to be tracked in the indicated sequences after departing Hanscom Field. The APT 

system is to be initialized by tracking, in sequence,

TABLE 5. Position coordinates and gravity values for retroreflector sites in the calibration range

Station

Name

Abbot -------------.

Acton -------------

Ashland -----------

Framingham -------

Haystack ----------

Indian -------------

Lab ---------------

Lincoln* -----------

Lake --------------

Littleton -----------

Mohawk -----------

Nagog -------------.

Nat ---------------.

Post ---------------.

Sudbury -----------

Water ..............

Latitudea

........ 42°35' 17.8072" N

........ 42°28'46.5206" N

........ 42°15' 6.7190" N

-------- 42°18' 9.6600" N

........ 42°37'21.4841" N

_.-_--_- 42°21' 0.5948" N

........ 42°27'52.5270" N

........ 42°28'23.5882" N

........ 42°40'59. 1785" N

-------- 42°32'42. 1234" N

........ 42°28'22.3676" N

........ 42°31'20.8391" N

........ 42°24'48.2871" N

........ 42°21'33.8249" N

........ 42°23'50.6621" N

........ 42°38'26.5968" N

Longitudea

71°25'57.6360" W

71°27'27.3357" W

71°28'29.7035" W

71°23'47.1111" W

71°29' 19.2579" W

71°30' 7.6261" W

71°17'58.8542" W

71°17'54.1945" W

71°23'20.4719" W

71°29' 6.4914" W

71°23'58.7138" W

71°26' 2.9862" W

71°29' 7.4322" W

71°25'37.2286" W

71°24' 0.7526" W

71°21'51.2507" W

Elevationb

(meters)

116.16

75.35

84.77

59.64

122.06

97.13

47.22

38.73

63.90

76.32

60.67

79.86

114.19

49.15

59.55

40.33

Gravity Value0

(milligals)

980,380.988

980,373.153

980,333.761

980,360.025

980,373.30

980,352.555

___

980,380.574

980,391.069

980,383.650

980,374.194

980,379.565

980,362.398

980,365.562

980,367.476

980,397.790

aPosition of retroreflector, referenced to the North American Datum of 1927. 

bElevation of retroreflector, referenced to the National Geodetic Vertical Datum of 1929.

 Measured at Survey bench mark tablet, and corrected for instrument drift and tidal effects.

*Control station only, without retroreflector.
88 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

the Mohawk, Nagog, and Littleton retroreflectors. 

Thereafter, four "counterclockwise" passes are to 

flown around the indicated (fig. 48) nearrectangular loop, a round-trip excursion to the Indian retroreflector, four "clockwise" passes around 

the foregoing loop, and return to Hanscom Field 

over the same retroreflectors (in reverse order) 

used to initialize the APT system.

Aircraft turns: All are to be made at standard rate.

Profiler and video camera: Operate while over the

Acton and Sudbury orthophoto map areas. 

Flight 4

Purpose: To determine accuracy of the APT system 

in terrain profiling, with tracker updates at about 

4-minute intervals.

Flight plan: Layout as shown in figure 49. Retroreflectors are to be tracked in the indicated sequences after departing Hanscom Field. The APT 

system is to be initialized by tracking first the 

Mohawk and then the Acton retroreflectors. 

Thereafter, four "clockwise" passes are to be flown 

around the indicated (fig. 49) loop, a round-trip

HAYSTACK

LITTLETON

WATER

INDIAN

ASHLAND

EXPLANATION

CONTROL POINT WITH 

RETROREFLECTOR

"OUTBOUND"TRACK 

DIRECTIONS ONLY

DISTANCE TRAVELED IN 1 

MINUTE AT 115 MPH 

(100 KNOTS)

0 5 MILES

5 KILOMETERS

ACTON

NAT

LAB (START AND END)

FRAMINGHAM

^

FIGURE 47. Flight plan for performance-evaluation flights 1 

and 2.

.LITTLETON

START

EXPLANATION

CONTROL POINT

WITH RETROREFLECTOR

ORTHOPHOTO MAP

'OUTBOUND" TRACK 

DIRECTION ONLY

DISTANCE TRAVELED IN 1 

MINUTE AT 115 MPH 

(100 KNOTS)

0 3 MILES

0123 KILOMETERS

FIGURE 48. Flight plan for performance-evaluation flight 3.

excursion to the Abbott retroreflector, four 

"counterclockwise" passes around the foregoing 

loop, and return to Hanscom Field over the same 

retroreflectors (in reverse order) used to initialize 

the APT system.

EXPLANATION

 CONTROL POINT

WITH RETROREFLECTOR

[~l ORTHOPHOTO MAP

-» "OUTBOUND"TRACK 

DIRECTION ONLY

DISTANCE TRAVELED IN 1 

MINUTE AT 115 MPH 

(100 KNOTS)

1 3 MILES

1 2 3 KILOMETERS

START AND END

FIGURE 49. Flight plan for performance-evaluation flight 4.
PERFORMANCE EVALUATION FLIGHTS 89

Aircraft turns: All are to be made at standard rate. 

Profiler and video camera: Operate while over the 

Nagog orthophoto map area.

FLIGHT RESULTS

In the following discussion, it will be apparent that 

some departures from the detailed flight specifications 

were necessary. These resulted from the recurring need 

to compromise with such disruptive factors as the 

threat of intolerable weather delays, unpredictable 

malfunctions of key instrument components, uncertainties in visual navigation, and proximity of other 

aircraft entering the calibration range airspace.

Flights 1 and 2

Although intended to be at least 1 day apart, both 

flights were accomplished on January 26,1984 one in 

the morning and one in the afternoon. All retroreflectors were tracked successfully in accordance with the 

specified order, flight altitude, and speed. The data 

were processed after the flights in a way that allowed 

surveying accuracy to be shown for three different 

spacings of control (retroreflectors). These spacings can 

be described, for a given flight line, in terms of the 

following retroreflectors for which it was assumed that 

position coordinates were known: 

a. Alternate (every other) retroreflectors. 

b. The two end and the midpoint retroreflectors. 

c. The two end retroreflectors only. 

In each of the foregoing cases, the position coordinates 

were computed from the flight data for all other 

retroreflectors. The values thus computed then were 

compared to the corresponding values determined in 

the original field surveys to ascertain the APT system 

surveying errors. Results of these two flights, and the 

subsequent data processing, are given in table 6. Parts 

A, B, and C of the table correspond to the foregoing 

retroreflector-spacing cases a, b, and c.

In part A, the position-coordinate errors are seen to 

range, in centimeters, from -23 to 38 for latitude, -31 to 

80 for longitude, and -27 to 10 for elevation. The corresponding standard errors, in centimeters, are ±25, 

±38, and ±13. Some of the error in the horizontal coordinates can be attributed to noise in the Doppler equipment used in the original field surveys. This is estimated to be ±10 to ±20 cm. The changes in errors, 

when the morning and afternoon flight data are compared, are seen to range, in centimeters, from -1 to 12 

for latitude, -43 to 15 for longitude, and -5 to 8 for 

elevation. The corresponding standard errors, in centimeters, are ±7, ±22, and ±6. Thus, the agreement 

between the morning and afternoon flight surveys is 

excellent and gives a good measure of the precision 

with which the APT system is performing; for example,

the smallness of the standard-error values indicates 

capability of the system to monitor, through periodic 

repeat flight surveys, land surfaces that are subject to 

long-term changes due to erosion or subsidence.

Inspection of the results given in parts B and C of 

table 6 shows that, even with substantial relaxation in 

the frequency with which the IMU is updated (through 

tracking of retroreflectors at known locations), overall 

accuracy and precision degrades little. Conversely, it 

can be concluded that even greater accuracy and precision might be achieved if the frequency of updates was 

increased through the use of additional retroreflectors.

Analysis of all data in table 6 supports the conclusion 

that overall performance capability for determining position and elevation coordinates will indeed be the 

specified ±60 and ±15 cm, respectively, 90 percent of 

the time, if the IMU is updated at 3-minute instead of 

5-minute intervals.

Flights 3 and 4

These flights were first flown, primarily as planned, 

on January 22, 23, 27, and 30,1984. However, frost and 

snow cover required almost daily clearing of the 

retroreflectors, and the fine-detail imagery (painted 

lines on parking lots, playgrounds, and tennis courts) 

needed to assess horizontal accuracy of the terrainprofile data generally was obscured. All retroreflectors 

were tracked successfully in accordance with the specified order, flight altitude, and speed. It was noted, however, that laser return signal dropouts occurred as the 

profiler passed over the few snow-cleared areas, which 

were commonly asphalt parking lots and roads. It was 

concluded that, at this flight altitude [2000 ft (610 m)], 

the black asphalt absorbed most of the laser energy. 

Thus, when all data were processed after the flights, 

the best terrain-profile points available for comparison 

with ground truth were those on the frozen surface of 

Nagog Pond and on the football fields at Sudbury and 

Acton High Schools, but even these comparisons suffered from the unknown vagaries of depth of snow 

cover.

Because of the foregoing uncertainties, flight 3 was 

repeated on May 10, 1984, and flight 4 on June 15, 

1984. Retroreflectors were tracked in accordance with 

the specified order and flight speed, but the flight altitude was lowered slightly to 1,650 ft (503 m) to minimize laser return-signal dropouts caused by asphalt 

surfaces. Results of only the May and June flights are 

given in table 7. The intent was simply to identify on 

each profile leg a few (four to eight) points for which a 

reliable and accurate elevation was interpolated on the 

orthophoto map. The average values for each set of 

points were then tabulated.

The average vertical error is the average difference 

between the profiler elevations and the orthophoto map
TABLE 6. Results of performance-evaluation flights 1 and 2

A. Alternate retroreflector positions known 

(5.2 min. between IMU updates)

Frror (cm) Change (cm)

Retroreflector

ABBOT

ACTON

MOHAWK

INDIAN

POST

Standard

Error

Retroreflector

ABBOT

NAGOG

NATICK

SUDBURY

INDIAN

POST

Standard

Error

Flight

am 

pm

am

pm

am

pm

pm

am

pm

B. End

Flight

am

pm

am

pm

am

pm

am

pm

am

pm

am

pm

Lat.

37 

38

- 1

- 7

-23

-23

34

28

- 1

-13

+25

Long.

X

-19 

-21

-13

-28

-22

-31

37

80

31

49

+38

Elev.

E A<() AX AE

-19 

-27 -128

3

- 4 6 15 7

- 2

3 0 9-5

-17

-12 6 -43 - 5

10

5 12 -18 5

+13 + 7 +22 + 6

C. End retroreflector positions kn

(15.7 min.

Retro- Lat.

reflector Flight <j>

ABBOT am 41

pm 73

NAGOG am 3

pm 37

ACTON am - 4

pm 20

MOHAWK am -22

pm 0

between IMU updates

Error (cm)

Long.

X

37

50

68

79

40

11

27

9

Elev.

E At

-12

-24 -3

8

2 -3

6

- 6 -2

2

6 -2

and midpoint retroreflector positions known

(7.8

Lat.

*

49

63

11

26

-58

-38

- 2

23

- 9

- 3

- 3

6

+32

min. between IMU updates)

Error (cm)

Long.

X

4

35

33

66

26

-72

5

-55

58

12

36

6

+41

Change ( cm)

Elev.

E A<() AX AE

-16

-23 -14 -31 7

4

3 -15 -33 1

-18

- 2 -20 98 -16

-14

1 -25 60 -15

-29

-16 - 6 46 -13

- 2

7 - 9 30 - 9

+14 +16 +55 +11

NATICK am -61

pm -23

SUDBURY am - 7

pm 26

INDIAN am - 9

pm 11

POST am - 5

pm 11

Standard

Error +30

(After Cyran and Chapman,

60

-65

15

-51

83

16

43

9

+48

1984) .

-11

-8 3

-15

2 -3

-23

-23 -2

- 1

7 -1

+12 +2
USES FOR THE INSTRUMENT SYSTEM 91

TABLE 7. Results of performance-evaluation flights 3 and 4

IMU 

Flight Orthophoto update 

date map site interva 

(min)

May 10, 1984 SUDBURY

ACTON

SUDBURY

ACTON

SUDBURY

ACTON

---do----

SUDBURY

ACTON

SUDBURY

ACTON

June 15, 1984 NAGOG

... do ---

---do ---

... do ---

... do ---

---do ---

... do ---

---do ---

6.6

6.4

7.4

6.3

7.5

8.0

4.8

7.0

4.8

7.1

4.9

4.6

5.0

5.3

3.9

3.9

9.3

9.3

3.8

Average Errors

1 Latitude

-30

90

-190

Longitude

-90

-40

20

No suitable point

-80

50

-40

30

-70

0

90

80

-60

0

80

-80

180

130

-200

-160

110

-70

50

60

No suitable point

No suitable point

-30

110

120

60

-100

-10

(cm)

Elevation

10

22

-3

7

-15

15

0

-20

16

-8

3

-2

11

7

-15

3

-15

1

4

elevations. The average horizontal errors are the average differences between the latitudes and longitudes as 

determined by the APT system and as scaled from the 

orthophoto map. The ranges in position coordinate errors, in centimeters, as shown in table 7, are from -20 

to 22 for elevation, -190 to 120 for latitude, and -200 to 

180 for longitude. The vertical errors are reasonably 

representative and tolerable; the horizontal errors are 

prejudiced to some degree by the small-scale video image and its lack of crisp definition.

Processing of all profiler data from all flights (January, May, and June) yielded a standard vertical error of 

±6 cm when the IMU is updated at 3-minute intervals. 

If this is combined with the noise value for the profiler 

itself of ±6 cm, then an overall performance capability 

results in an elevation position coordinate of ± 14 cm 90 

percent of the time, which is within the ±15 cm given 

in the specifications for the APT system.

USES FOR THE INSTRUMENT SYSTEM

The stimuli for this instrument-development project 

have persistently arisen from field-investigational 

problems for which more precise, faster, and less costly 

(with less manpower) data-gathering techniques are 

perennially sought. The list of potential uses for the 

instrument system appears endless as field programs of 

the many agencies engaged in earth science studies are 

reviewed. In the following sections, descriptions are 

given for a few selected uses drawn from the more obvious and pressing requirements in field programs of the 

Geological Survey.

HYDROLOGIC STUDIES

As part of its Congressional mandate to evaluate the 

Nation's water resources, the Survey continuously 

maintains a nationwide network of stream-gaging stations, sediment-sampling stations, ground-water-level 

observation wells, and water-quality sampling and observation stations. The network is in constant selective 

revision as some stations are discontinued and new 

ones are introduced.

Although the data regularly collected from the foregoing station networks are used for many purposes, one 

particular use is of special interest here. That use relates to the large variety of mathematical models being 

designed to describe and predict such everyday phenomena as surface runoff from watersheds of many 

sizes and types, which include pristine as well as manaltered states (examples of the latter are strip-mined or 

urbanized watersheds); open-channel flow characteristics in selected reaches of streams in various geologic 

settings; the spread and migration of contaminants introduced in various ways into the flow regimes in 

watersheds and stream channels; the land subsidence 

in and around heavily pumped ground-water reservoirs 

and producing geothermal reservoirs; and erosional 

processes along beaches and shorelines.

In any given modeling problem, the data from the 

nearest available network station or stations constitute only a small fraction of the total data requirement. 

A large percentage of what is lacking can best be described as precise geometric details (three-dimensional 

shape) on the surrounding land surface or stream channel or beach or shoreline. Commonly, such details cannot be obtained with the needed precision from any 

available maps. Thus, recourse, heretofore, could only 

be through arrangements for special surveys of the desired land or channel area using field parties and conventional surveying or mapping techniques.

The advent of the APT system brings into focus the 

following immediate and specific uses related solely to 

ongoing nationwide hydrologic studies:

1. To measure stream-valley cross sections at V^mi 

(0.4-km) spacings, along all streams that are represented in the station networks.

2. To measure stream-valley cross sections at designated spacings along stream reaches that are encompassed in specific one- and two-dimensional 

mathematical models that are to be used for describing open-channel flow.

3. To measure terrain profiles at designated sites in 

watershed areas modified by man that are encompassed in specific mathematical models which are 

to be used for describing changes in surface 

runoff.

4. To measure precise positions for ground-control 

points at designated locations in areas where ob-
92 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

servation-well positions in a local network must 

be accurately fixed.

5. To measure terrain profiles periodically at designated sites in areas where innumerable natural 

ponds are expressions of the local water table. 

These data are to augment and enhance the conventional water-level observation-well data and 

will allow preparation of more detailed watertable contour maps for use by water managers.

6. To measure precise positions periodically for 

ground-control points at designated locations in 

subsidence areas. Initially, the area choices will 

be limited arbitrarily to those that have a maximum diameter of less than 10 mi (16 km).

7. To measure longitudinal profiles of streams in flood, 

as opportunities occur, to enhance and extend the 

conventional flood-stage data that commonly are 

collected only at selected and, perhaps widely 

spaced, points.

MAP PRODUCTION TASKS

The National Mapping Division conducts the National Mapping Program, which provides graphic and 

digital cartographic products and information for the 

United States, Territories, and U.S. possessions. The 

products include several series of topographic maps, 

land use and land cover maps and associated data, geographic names information, geodetic control data, and 

remote sensing data. The 7V2-minute 1:24,000- and 

l:25,000-scale topographic maps in the conterminous 

United States and Hawaii and the 15-minute 1:63,360- 

scale maps in Alaska provide basic coverage of these 

areas and are the building blocks used for developing 

almost all other cartographic products. Over 46,000 

maps in the series have been completed, and the total 

national coverage of about 55,000 maps is planned for 

this decade. Present work includes new mapping to 

complete the coverage and maintenance of the existing 

maps. The APT system can provide support to this program in the following ways:

1. In the construction of topographic maps, map details 

and contour lines are compiled using stereoscopic 

models, which are formed by merging pairs of 

aerial photographs. Basic control to orient, scale, 

and position the models generally is acquired by 

field survey and aerotriangulation methods. The 

APT system can provide model control by establishing terrain profiles in the side-lap areas midway between the aerial photograph flight lines. 

The APT system video imagery provides the connection to the aerial photography for control point 

identification and selection.

2. Terrain-profile data also can assist in the map compilation process by providing ground elevations in

areas of tall timber. The operator of a photogrammetric instrument seldom sees the ground under 

tall timber conditions and must estimate tree 

heights to position the contours. The APT system 

profiles can provide a sampling of tree height 

data, as well as elevations, so that contour estimates can be more accurate.

3. Some maps may not meet National Map Accuracy 

Standards. From 1944 to 1958, the Survey produced Ti/^-minute maps using photogrammetric 

methods that were in their early phases of development. In some instances, the technology and 

methodology may not have produced maps that 

meet the high-accuracy standards of today. Many 

of the maps produced during this period are now 

candidates for revision or replacement. The APT 

system has the capability to determine the absolute accuracy of the maps in question and, based 

on these findings, to assist in determining the 

most cost-effective means of updating. Accuracy 

testing will consist of a series of terrain profiles 

spaced across the map and following along roads 

and other linear features. Points common to the 

map and the APT video system will be selected, 

and positions and elevations from both sources 

obtained. Over 20 points will need to be compared 

to determine the horizontal accuracy of the map 

and another 20 points to determine the vertical 

accuracy.

4. The video imagery as recorded by the APT system 

has a potential application in the map revision 

program. The images will be collected at a flight 

height of several thousand feet over an area to be 

revised. The images can be computer processed to 

correct for tip, tilt, and scale; then oriented to 

north, and displayed as an overlay to the map 

image. New buildings, roads, reservoirs, shopping 

centers, and other features then can be identified 

quickly and captured for placement on the revised 

map.

The APT system has the capability of developing into 

a versatile mapping tool. The video image provides a 

record of the planimetry, and suitable scale can be selected by varying the flight height or changing the 

camera zoom setting. Because the image will be digitally processed, geometric corrections can be made to 

account for variations in flight height and orientation. 

The profile data provide ground elevations that can be 

used to develop contours or digital elevation models. If 

a scanning capability ultimately is added to the profiler, a swath of elevation data can be obtained in a 

single pass. The area to be mapped can be covered with 

many parallel or, for more accuracy, crisscrossing 

flight lines. Flight-line spacing, as well as flight 

height, can be changed to meet the specified scale and
USES FOR THE INSTRUMENT SYSTEM 93

accuracy of the desired map product. Because the video 

image and the elevation data will be processed in a 

computer, it appears that further research in this direction will produce a nearly automatic mapping system.

GEOLOGIC STUDIES

The APT system has the capability for improving the 

quality of routine airborne or ground-based surveys 

and also can serve as an experimental device from 

which three-component, gravity-disturbance information can be extracted.

Routine airborne surveys include those that use electromagnetic, gas or vapor detection, magnetic, photographic, and radioactive techniques. All have a critical 

dependence upon accurate knowledge of aircraft position as a function of time of measurement. The APT 

system can determine aircraft position more accurately 

than existing radar or radio transponders.

Routine ground-based surveys include those designed for measuring gravity anomalies and determining the longitudinal profiles of stream valleys. Both 

types of survey are critically dependent on the accuracy 

of elevation (vertical coordinate) control. The APT system can determine the vertical coordinate to within 

6 in. (0.15 m) and, thereby, offers elevation control sufficient for gravity-meter profiling or for correlating 

changes in stream gradient with geologic contacts between rock types of contrasting erosional resistance.

Extraction of three-component, gravity-disturbance 

information from the APT system depends, in part, on 

future contract work to refabricate the system more 

closely in accordance with the original design specifications. In that format, the system offers promise as an 

airborne gravity mapper, with a capability for achieving a gravity-anomaly accuracy to within 0.2 milligal, 

and vertical-deflection accuracy to within 0.1 arcsecond. An airborne gravity mapper suffers, relative to 

land-based gravimetry, from attenuation of highamplitude, short-wavelength anomalies that relate to 

locally concentrated, shallow-depth sources. However, 

such a mapper benefits enormously from the standpoints of unlimited accessibility, speed and density of 

coverage, and spatial consistency of all simultaneously 

measured data. The principal use of the APT system as 

a gravity mapper is envisioned for nonrepeat surveys 

over terrains of low-to-moderate relief characterized by 

low-to-moderate gravity gradients.

Economic feasibility eventually will prompt the development of the next and even more accurate generation of the APT system. When this occurs, the system 

can be applied more widely to the important realm of 

precise repeat surveys that involve the measurement of 

differential ground movements over selected time 

intervals. Such repeat surveys are vital to a variety of 

geologic studies that include monitoring of active soil

creep, movement along major strike-slip fault zones, 

and landsliding; volcano inflation and deflation; ongoing beach and slope erosion; the melting and growth of 

glaciers; and land subsidence in and around producing 

oil fields, geothermal reservoirs, and heavily pumped 

ground-water reservoirs.

FUTURE IMPROVEMENTS IN THE 

INSTRUMENT SYSTEM

Some clues already have been given on highly desired uses for the APT system when it can be built to 

conform as precisely as possible to the full set of original design specifications. Important elements in those 

specifications concern the need to minimize overall system size, weight, and power requirements so that the 

logistics for and cost effectiveness of regular field operation will be minimized and maximized, respectively. 

Of equal importance is the need to provide a stable 

platform for the profiler slaved to the stable platform of 

the inertial navigator. The present technology in all 

pertinent scientific fields is advanced beyond the points 

needed to meet the original specifications. Thus, an 

efficient field-operational APT system can be built 

whenever the administrative decision dictates. Such a 

step is one viable way of improving the present system.

Other improvements are foreseeable, however, that 

would allow the original specifications to be refined in 

several desirable ways. Three leading-candidate improvements, out of a longer list of possibilities, relate to 

a scanning-laser device, a gravity-gradiometer device, 

and a global positioning system (GPS) receiver.

An expanding variety of engineering and environmental problems has created a large demand for digital 

models of terrain that can then be analyzed by computer. As an adjunct to the APT system, a laser device 

capable of side-to-side scanning in a direction across 

the aircraft track would provide a wealth of terrain 

elevation data for digital-modeling purposes. Furthermore, the accuracy of such data would greatly exceed 

that obtainable through such existing techniques as 

digitizing available topographic maps or application of 

photogrammetric methods. Ideally, a large scan angle 

would maximize the amount of data obtained along the 

flight path; practically, the scan angle is usually limited to about 10° or 15° either side of the vertical to 

obtain reasonable penetration of vegetation.

In the Federal establishment, interest is building for 

the development of an airborne gravity gradiometer. 

Over the past decade, the Navy Department and the 

Air Force have supported developmental work by such 

contractors as Draper Laboratory, Bell AerospaceTextron Laboratories, and Hughes Research Laboratories. Working models should become available in the
94 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

near future, and one such device would be an ideal 

companion for the present and the improved versions of 

the APT system. The combination would permit simultaneous extraction of the highest quality gravity data 

and highest quality positional data yet available from 

an aircraft. This would immediately extend aerial surveying operations to terrains of higher relief, where 

gravity-anomaly gradients commonly are severe and 

the spacing between ground-surveyed retroreflectors 

may need to be increased greatly because of difficult 

ground access.

The idea of combining a GPS receiver with the APT 

system in the same aircraft is especially attractive because it appears that the individual position-error models for the two instruments will complement each 

other. This means that a given flight mission could be 

flown with significantly fewer retroreflectors without 

any degradation in accuracy. An added benefit would 

accrue from the much slower growth in the real-time 

navigation error, which would ensure greater success 

in tracker search for and lock-on to each retroreflector 

programmed for use during the mission.
SELECTED REFERENCES 95

SELECTED REFERENCES

Ayres, Frank, Jr., 1954, Plane and spherical trigonometry: 

New York, McGraw-Hill, 207 p.

Birkhoff, Garrett, and Bartee, T. C., 1970, Modern applied 

algebra: New York, McGraw-Hill, 431 p.

Britting, K. R., 1971, Inertial navigation systems analysis: 

New York, John Wiley, 249 p.

Brooks, F. P., Jr., 1975, The mythical man month Essays on 

software engineering: Reading, Mass., Addison-Wesley 

Publishing Co., 195 p.

Broxmeyer, Charles, 1964, Inertial navigation systems: New 

York, McGraw-Hill, 254 p.

Chapman, W. H., and Brown, R. H., 1977, Applications of 

inertial technology in Geological Survey programs, in 

First International Symposium on Application of Inertial 

Technology to Surveying and Geodesy, Ottawa, Canada, 

October 12-14, 1977, Proceedings: p. 222-229.

Chapman, W. H., and Starr, L. E., 1979, Surveying from the 

air using inertial technology: Annual Convention 

ACSM/ASP, Washington, B.C., March 1979, Proceedings, p. 352-365.

Cyran, E. J. and Chapman, W. H., 1984, Flight testing the 

aerial profiling of terrain system (APTS): U.S. Geological 

Survey Open-File Report 84-354, 18 p.

Desai, Mukund, Drohan, W. A., Hursh, J. W., Mamon, Glenn, 

and Youmans, D. G., 1975, Aerial profiling of terrain to 

define stream-valley geometry Study report: The 

Charles Stark Draper Laboratory, Inc., Cambridge, 

Mass., Report R-870, 225 p. (Also, U.S. Geological Survey Open-File Report 76-672.)

Draper, C. S., Wrigley, Walter, and Hovorka, John, 1960, 

Inertial guidance: New York, Pergamon Press, 130 p.

Forsythe, A. I., Keenan, T. A., Organick, E. I., and Stenberg, 

Warren, 1969, Computer science A first course: New 

York, John Wiley, 553 p.

Hill, F. J., and Peterson, G. R., [1978], Digital systems: Hardware organization and design (2nd ed.): New York, John 

Wiley, 701 p.

Hosmer, G. L., [1948], Practical astronomy (4th ed.): New 

York, John Wiley, 355 p.

Hursh, J. W., Mamon, Glenn, and Soltz, J. A., 1977, Aerial 

profiling of terrain, in First International Symposium on 

Application of Inertial Technology to Surveying and 

Geodesy, Ottawa, Canada, October 12-14, 1977, Proceedings: p. 122-130.

Kalman, R. E., 1960, A new approach to linear filtering and 

prediction problems: American Society Mechanical Engineers Transactions, Series D, Journal of Basic Engineering, v. 82, p. 35-45.

Landau, L. D., and Lifshitz, E. M., [1976], Mechanics (3d ed.): 

New York, Pergamon Press, 169 p.

Lapedes, D. N., ed., [1978], Dictionary of scientific and technical terms (2nd ed.): New York, McGraw-Hill, 1771 p.

Mamon, Glenn, Wheeler, R. L., and Youmans, D. G., 1976, 

Laser ranging and tracking system analysis and experiments for aerial profiling of terrain: Cambridge, Mass., 

The Charles Stark Draper Laboratory, Inc., Report R968, 51 p.

Mamon, Glenn, Youmans, D. G., Sztankay, Z. G., and Mongan, C. E., 1978, Pulsed GaAs laser terrain profiler: Applied Optics, v. 17, no. 6, p. 868-877.

Pedde, L. D., and others, 1978, Metric manual: U.S. Department of the Interior, Bureau of Reclamation, 278 p.

Pitman, G. R., Jr., ed., 1962, Inertial guidance: New York, 

John Wiley, 481 p.

Soltz, J. A., Mamon, Glenn, and Chapman, W. C., 1981, 

Aerial profiling of terrain system implementation, in 

Second International Symposium on Inertial Technology for Surveying and Geodesy, Banff, Alberta, Canada, 

June 1981, Proceedings: The Charles Stark Draper Laboratory, Inc., Cambridge, Mass., Report P-1343, 23 p.

The Charles Stark Draper Laboratory, Inc., 1977, Aerial profiling of terrain system design-phase interim report, R1070: Cambridge, Mass., open-file report, 401 p.

 1978, Aerial profiling of terrain system design-phase 

final report, R-1219: Cambridge, Mass., open-file report, 

372 p. 

-1982, Aerial profiling of terrain system, phase V fabrication, assembly, and testing, R-1451: Cambridge, 

Mass., open-file report, 2632 p.

U.S. Department of Commerce, 1977, The international system of units (SI): National Bureau of Standards Special 

Publication 330, 41 p.

Vanderwall, Jonathan, Hattery, W. V., and Sztankay, Z. G., 

1974, Subnanosecond rise time pulses from injection 

lasers: IEEE Journal of Quantum Electronics, v. 10, 

no. 7, p. 570-572.

Youmans, D. G., 1977, Flight testing of an airborne laser 

terrain profiler: The Charles Stark Draper Laboratory, 

Inc., Cambridge, Mass., Report R 1106, 36 p.
96 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

SYMBOLS AND DIMENSIONS*

Symbol

kg

m

s

N

G

M

k

m

Mn

P

rad

Hz

Pa

T

O)

L

V

Conventional

Units Dimensions Description

kg M kilogram, the unit of

mass

m L meter, the unit of

length

s T second, the unit of

time

m-kg-s~2 LMT~2 newton, the unit of

force

(1) giga, a multiplying

factor of 1 x 109

(1) mega, a multiplying

factor of 1 x 106

(1) kilo, a multiplying

factor of 1 x 103

(1) milli, a multiplying

factor of 1 x 10"3

(1) micro, a multiplying

factor of 1 x lO'6

(1) nano, a multiplying

factor of 1 x 10"9

(1) pico, a multiplying

factor of 1 x 10- 12

rad (1) radian, the unit of

measure for a

plane angle (360°

= 2ir radians). It

is the plane angle

between two radii

of a circle which

cut off a circumferential arc

whose length

equals the radius.

S" 1 T" 1 hertz, the unit of

frequency, usually

expressed as number of cycles per

second

N-m- 2 L" 1MT- 2 pascal, the unit of

pressure, usually

expressed in newtons per square

meter

N-m or L2MT" 2 torque, the moment

m2 kg s " 2 of force, usually

expressed in newton meters

rad-s" 1 T" 1 angular velocity,

usually expressed

in radians per second

N-m-s or L2MT-1 angular momentum,

m2 -kg-s" 1 usually expressed

in newton meter

seconds

m-s" 1 LT" 1 velocity, usually expressed in meters

per second

Conventional

Symbol Units Dimensions Description

g m-s" 2 LT" 2 acceleration due to

gravity, usually

expressed in meters per second

squared. The

standard value is

_^ 9.80665

7l s-1 T" 1 Earth rate of rotation or angular

velocity vector,

usually expressed

in radians per second

RM m L Earth radius of curvature in the meridian plane, usually expressed in

meters

RN m L Earth radius of curvature normal to

the meridian

plane, usually expressed in meters

O latitude, the angular

distance, in degrees, of a point

on the Earth's

surface north or

south of the equator, measured

along a meridian

(the equator is lat.

0°, the North Pole

lat. 90°N, and the

South Pole lat.

90°S)

\ longitude, the angular distance, in degrees, west or

east, measured

along the equator

or a parallel of

latitude, between

the meridian

through a point on

the Earth's surface and the reference meridian

(long. 0°) through

Greenwich, England (longitude

values increase to

a maximum of

180°W or 180°E)

x m L position coordinate

(north-south direction) in the horizontal (locally

level) plane
SYMBOLS AND DIMENSIONS 97

Symbol

Conventional 

Units Dimensions Description

y m L position coordinate

(east-west direction) in the horizontal (locally 

level) plane

z m L position coordinate

(local vertical direction) in reference to the NGVD 

of 1929

a rad (1) angle of divergence,

usually expressed 

in radians

bar N-m~2 L-1MT~2 bar, a unit of pressure, usually restricted to meteorology, equal to 

105 pascals

cd candela, the base

unit of luminous 

intensity (optics), 

formerly known as 

candle and new 

candle

sr steradian, a unit

solid angle with 

its vertex at the 

center of a sphere, 

and of a size such 

that it subtends a 

spherical surface 

area equal to a 

square, each of 

whose sides equals 

the sphere radius

1m cd-sr lumen, the unit luminous flux, 

which is the flux 

emitted within a 

unit solid angle (1 

steradian) from a 

point source having an intensity of 

1 candela

Ix lm-m~2 or lux, the unit of illucd-sr-m~2 mination also

called the illuminance, which is 

the illumination 

on a surface area 

of 1 square meter, 

over which there 

is a uniformly distributed luminous 

flux of 1 lumen

"International System of Units (SI). 

(l)Dimensionless.
INDEX

(Italic page numbers indicate major references)

Acceleration. See Aircraft motion, Inertia! navigator.

Accelerometer _________ 3, 4, 6,7, 9,10, 15-19, 33, 34 

Array _______________________ 15, 18, 19 

Closed-loop system ___________________ 16 

Electromagnetic suspension _______________ 16 

Floated gyro ________________________ 16 

Gravity measurement _____ 27, 29, 34, 75, 76, 87, 93, 94 

Input axis _______________________ 16-19 

Life expectancy ______________________ 18 

Measurement error __________________ 35, 36 

Pendulosity _______________________ 18 

Pendulous mass ______________________ 16 

Resolver __________________________ 17 

Sensitive axis _______________________ 18 

Sensitivity _________________________ 18 

Signal generator _____________________ 16 

Torque ________________________ 15-17 

Velocity measurement _________________ 33,83

Accuracy. See Horizontal or Vertical accuracy.

Aerial calibrate. See Modes, APT system.

Aerial survey. See Modes, APT system.

Aerial profiling of terrain system __ 1, 4, 7-9, 27, 28, 34, 36,

43, 46, 49, 50 , 55, 56, 70, 73, 74, 76-81, 93, 94

Maintenance ________________________ 84

Performance evaluation _______________ 85-91

Uses ________________________ 9, 91-93

Airborne computer. See Computer, digital, flight.

Airborne gravity gradiometer _____________ 93, 94

Aircraft ______________________ 2, 3, 6, 76-78 

Flight path ___________ 32, 35-37, 78, 81, 83, 84, 93 

Motion ______________ 7,19, 24, 28, 29, 33-35, 43

See also Pitch motion, Roll motion, Yaw motion. 

Operating altitude _____________________ 45 

Speed __________________________ 81 

Turning rate ________________________ 81

Air loop. See Temperature control.

Alarm ________________________ 51, 58, 60

Altimeter, aircraft. See Laser tracker, pressure transducer.

Angular momentum. See Gyro.

Angular rotation. See Aircraft motion, Coordinate-axis system.

Angular velocity. See Earth rotation, Gyro.

Anisoelasticity ________________________ 14

Antireflection coating ____________________ 46

APT system. See Aerial profiling of terrain system.

Astronomical latitude. See latitude coordinate.

Axis

Azimuth ________________________ 19, 24 

Coordinate. See Cartesian coordinates, Coordinate-axis

system

Elevation _______________________ 19, 24 

Input. See Accelerometer, Gyro.

Orthogonal ___________________ 7,11, 29-31 

Output. See Gyro.

Polar __________________________ 29, 31 

Roll __________________________ 19, 24 

98

Axis-Continued

Sensitive. See Accelerometer, Gyro.

x, y, z. See Cartesian coordinates, Coordinate-axis system. 

Azimuth motion. See Yaw motion. 

Azimuth stabilization loop 24

B

Back filtering. See Data processing.

Ball bearings. See Gimbal.

Baroaltimeter. See Laser tracker, pressure transducer.

Base rotation. See Gyro.

Beam splitter. See Laser tracker.

Bellows. See Gyro.

Bench marks _ 80, 81

Beryllium ________ 14

Buffer zone. See Temperature control.

Calibrate and aline. See Modes, APT system.

Calibration range ___ _ 85-87

Cartesian coordinates __________ - 12,15,18

See also Coordinate-axis system. 

Clock, master. See Master quartz-crystal oscillator. 

Closed-loop system. See Accelerometer, Inertial navigator. 

Color TV camera _______________ 59, 73, 78, 88, 89 

Computer, digital, flight ___________ 8,33-35, 46-66

Address ______ 50 

Cache ___________________________ 50 

Central processor ____ __ 49, 50, 60 

Channel _______________________ 49, 50-52 

Control panel _______ ___ 50 

Data processing ____ ___ 59, 60-66

Post flight ____________________59, 78, 79 

See also Data processing, Postmission.

Real time ___________________ 10, 50, 51, 60 

Direct memory access (DMA) ___ _ _ 51 

Double precision ___ _ 50 

Floating-point processor _____ _ _ 50 

Line printer ______________ _ 52, 53 

Magnetic tape cartridge _ ___ _ 54 

Magnetic tape controller _ ____ _ 53, 54 

Magnetic tape recorder _ __ ___ 53, 54 

Master-timing signal. See Time-code signal, Master quartzcrystal oscillator.

Memory ________________ ___ 49, 50 

Norden, PDP-11/70M _________________ 47, 49 

Packaging _______________________ 47, 48 

Parity check _______________ ___ 50 

Peripherals. See specific item under Computer. 

Port, input-output _________________ 49, 50-52 

Priority level ___________________ 49, 60, 61 

Program. See Computer software.

Register ________________________ 49, 50 

Software ________________________ 54-66
INDEX 99

Computer, digital, flight-Continued 

Software- Continued

Definition ________________________ 54 

Executive program _______________ 56, 59, 60 

Input-output driver routines ______________ 60 

Kalman filtering ___________________ 57, 85 

Master operating program ____________ 56, 60, 61 

Modes, system ____________________ 56-59

See also Modes, APT system.

Programs ___________________54-56, 59, 60

Structure ______________________ 59, 60

Table-driven programs _________________ 55

Tasks _________________________ 60-66

Asynchronous ___________________60, 61

Synchronous ____________________ 60, 61

Speed ____________________________ 49

Terminal, operator's ____________________ 51

Keyboard _________________ 25, 51, 52, 56, 57

Plasma-display panel ______________ 51, 56, 57

Throughput rate ______________________ 49

UNIBUS ______________________ 49, 50, 53

Constant-fraction discriminator. See Laser tracker.

Control. See Horizontal or Vertical control.

Convergence of meridians ________________ 29-31

Cooling coils _________________________ 27

Coordinate-axis system __ 3, 4, 7, 9-11, 19, 32, 35, 36, 78, 79,

83, 84

Angular rotation __________________ 29-31, 32

Positive sense ______________________ 31

Right-hand thread rule _________________ 31

Elevation coordinate __________ 9, 32-36, 78, 79, 93

Latitude coordinate ___________ 9, 29-33, 36, 78, 79

Longitude coordinate __________ 9, 29-33, 36, 78, 79

See also Cartesian coordinates.

Coriolis acceleration ___________________ 34, 35 

Corner-cube reflector. See also Retroreflector. __ 45, 46, 78,

79, 81

Cost-effectiveness study ____________________ 4 

Couple ____________________________ 11

D

Data format __________________________ 84 

Data processing. See Computer, digital, flight.

Back filtering _____________________ _ 85

Compressing _____________________ 84, 85

Editing __________________________ 84

Optimal smoothing ____________________ 85

Postmission __________________ 79, 83, 84, 85

Recombining _______________________ 85

DeHavilland Twin Otter aircraft. See Aircraft.

Delay line ___________________________ 73

Design phase ________________________ 4, 5

Design review _________________________ 5

Design verification phase __________________ 5, 6

Deviation, feedback loop _________________ 23, 24

Deviation sensor _______________________ 23

Deviation signal ______________________ 23, 24

Digital computer. See Computer, digital, flight.

Digital-to-analog channel. See Computer, channel.

Divergence angle. See Laser profiler, Laser tracker. 

Doppler satellite relative positioning technique ____ 86, 89 

Drift. See Gyro, Inertia! navigator.

E

Earth

Angular velocity ____________________ 31, 32 

Ellipsoidal shape ____________________ 29, 32 

Radius of curvature ____________________ 33 

Rotation, vector _______________ 7, 14, 29, 31, 32

Editing. See Data processing.

Electrical heaters. See Temperature control.

Electric oven _________________________ 43

Electromagnetic suspension. See Accelerometer, Gyro.

Electronics ________________________ 24, 25

Elevation coordinate. See Coordinate-axis system.

Elevation gimbal _________________ 19, 24,

Elevation stabilization loop _______________.

Ellipsoid __________________.._____ 9, 

See also Earth, Reference ellipsoid.

Encoders, resolver ____________________.

Engineering analysis phase _______________.

Equator __________________.___________

Equatorial plane ________________________________

Executive program. See Computer, software.

Eye safety ________________________ 40, 70

26,27 

 24

29,75

___24 

__4 

29,31 

29,30

Fabrication phase _______________________ 6 

Fans

Constant speed ____________________ 26, 27 

Variable speed ______________________ 27 

Feedback loop ____________________ 16, 23, 24 

Filter, optical _______________________ 38, 70 

Flight altitude ________________ 45, 81, 87, 89, 92 

Flight data ________________________ 84, 89 

Flight mission ________________ 81, 83, 84, 93, 94 

Flight path. See Aircraft, flight path.

Flight plan __________________ 59. 81, 83, 87-89 

Float. See Accelerometer, Gyro. 

Floated gyro. See Accelerometer, Gyro. 

Flood-plain mapping ____________________ 3. 81 

Footprint, laser ____________________ 41, 70, 71 

Frequencies, timing _____________________ 43 

Frequency stability ______________________ 43 

Friction torque ________________________ 19

G

Gain-control loop _______________________ 38 

Gallium-arsenide transmitter ____________ 38, 40, 70 

Geodesy __________________________ 29, 31 

Geodetic Survey of Canada __________________ 5 

Geographic pole ______________________ 29, 30 

Gimbal 

Angle measurement. See Inertial measurement unit, Laser

tracker.

Angular rate ________________________ 11 

Axis __________________________ 10-12
100 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

Gimbal - Continued

Ball bearings ________________________ 19

Elevation ____________________ 19, 24, 26, 27

Frequency, resonant _________________.__ 19

Friction torque ____________________.__ 19

Lock _____________________________ 36

Roll _____________________________ 19, 24

Servomotor. See Gimbal, torque motor.

Structure, IMU __________________ 7, 9-12, 19

Structure, tracker ___________________ 36, 37

Torque motor, IMU __________________ 19, 24

Torque motor, tracker _______________ 38, 41, 43

Glass window _______________________ 45, 46

Global positioning system (GPS) receiver ________ 93, 94

Gradiometer. See Airborne gravity gradiometer.

Gravity anomaly ___________________ 32, 93, 94

Gravity disturbance _________________ 32, 85, 93

Gravity-field vector _______ 7, 9,15,19, 28, 29, 33, 34, 36

Gravity gradiometer. See Airborne gravity gradiometer.

Gravity measurement ________________ 36, 87, 93

See also Accelerometer, gravity measurement. 

Gravity meter. See LaCoste and Romberg geodetic gravity

meter.

Gravity model ______________________ 1, 9, 83 

Ground-control point _ 3, 4, 7, 36, 37, 45, 77-79, 81, 83, 85-87,

91, 92

Gyro ____________________ 4, 6, 7, 9, 10-15, 32 

Angular velocity ____________________._ 14 

Array __________________________ 14, 15 

Base rotation ______________________ 11, 12 

Bellows ___________________________ 14 

Bias _____________________________ 34 

Center-flange mount ____________________ 15 

Damping fluid _____________________ 12, 14 

Drift _______________________ 7, 14, 34, 36 

Electromagnetic suspension ________________ 12 

Equation ________________________ 11,14 

Feedback loop _____________________ 23, 24 

Float ________________________ 12, 14, 15 

Flotation temperature ___________________ 12 

Gas bearing ______________________ 12, 14 

Input axis __________________ 11,12, 14, 15, 23 

Null ___________________________ 23, 24 

Output axis _____________________ 11, 14, 15 

Physical properties ____________________ 12 

Precession ___________________ 11, 16, 17, 32 

Precession torque _____________ 11, 12, 16, 17, 33 

Rotor __________________________ 10, 12 

Rotor speed ______________________ 11, 12 

Sensitive axis. See Gyro, input axis.

Sensitivity _________________________ 14 

Shims, thermal _______________________ 12 

Signal generator ____________________ 11, 12 

Single-degree-of-freedom _________________ 12 

Spin axis ______________________ 10, 11, 14 

Spin velocity ______________________ 11,12 

Time characteristic ____________________ 14 

Torque _____________________ 11, 14-16, 32 

Torque generator ______________________ 12

Gyro - Continued

Viscous shear _ - 14 

Wheel ________________ 10-12 

Gyrocompassing 33 

Gyroscope. See Gyro.

H

Heat dissipation. See Temperature control.

Heat exchanger. See Temperature control.

Hemisphere, northern 29, 32

High-frequency motion 7, 8

See also Aircraft motion.

History of work _____ 3-6 

Hooke's Law ___ 15 

Horizontal accuracy ___ 1, 89-92, 

Horizontal control _ 1, 77-79 

Horizontal Coordinates __ _ 32-34

See also Cartesian coordinates, Cordinate-axis system,

latitude and longitude coordinates.

Horizontal plane __ 12, 19, 27, 33 

Hydrogen-gas bearing 12

See also Gyro, gas bearing.

I

IMU. See Inertial measurement unit.

Inertia __________________ 9, 15

Inertial guidance _____ _ 2, 3, 9

Inertial measurement unit _ 7, 8, 9, 10, 27

Acceleration ____ 8

See also Accelerometer.

Angle measurement _ _ 83 

Case __________________________ 19, 38 

Laboratory calibration ______ _ 34 

Locally level type ____________ __ 9 

Stable platform _ 2, 7, 9, 10, U, 15, 19, 20, 23-27, 29, 31-34,

36, 79, 93

See also Gyro, Temperature control.

Tangent-plane type _______________ 9, 27, 29, 33 

Intertial navigator 1-4, 7, 8, 9, 10, 27, 28, 29, 35, 36, 78, 79, 93 

Acceleration _____________________ 8, 28, 29 

Aerial calibration ____________ 36, 59 

Azimuth alinement ___________ _ 33 

Calibration and alinement ________ _ 54, 57, 58 

Closed-loop system ___________________33-35 

Dead reckoning ______________________ 28 

Drift _____________________________ 7 

Gyrocompassing ______________________ 33 

Initial conditions ___________________ 28, 32, 33 

Movement

Earth rotation ___________________ 29, 31, 32

Eastward or westward _______________ 29-31

Northward or southward ______________ 29, 31

Navigation _________________________ 58

Orientation ____________________ 7, 29, 32, 33

Position ___________________________ 7, 83

Self alinement _____________________ 29, 33

Update _________________ 4, 7, 35, 36, 83, 87-91

Input axis. See Accelerometer, Gyro.
INDEX 101

International System of units ______________ 11, 66

Jitter _____________________________ 23

K

Kalman filter _______________________ 57, 85 

Keyboard _____________________ 25, 51, 56, 57

LaCoste and Romberg geodetic gravity meter _____ 87

Laser profiler _____________2-5, 36, 38, 40, 41, 66-73

Delay line __________________________ 73

Distance measurement __________________ 66

See also Laser profiler, range measurement. 

Divergence angle _____________ _ 70 

Eye safety _________________________ 70 

Filter, optical ________________________ 70 

Footprint __________________________ 71 

Last-pulse discriminator __________________ 73 

Mount structure ____________________ 73, 78 

Optics ____________________________ 70 

Power, peak ________________________ 70 

Precision of measurement _________________ 66 

Pulse rate __________________________ 70 

Range equation ____________________ 70, 71 

Range, maximum ______________________ 70 

Range measurement ____________ 70, 71, 73, 78, 84 

Signal processor __________________ __ 72 

Signal-to-noise ratio ____________________ 71 

Time-interval counter ________________ 43, 72 

Transmitter ________________________ 70 

Velocity of light ______________________ 66 

Wave form _________________________ 70 

Wavelength ________________________ 70 

Laser tracker ________________ 4, 7, 8, 35, 36-46 

Angle measurement ______________ 38, 41, 43, 84 

Beam path _______________________ 38, 41 

Beam splitter ________________________ 38 

Constant-fraction discriminator ______________ 43 

Divergence angle ______________________ 40 

Eye safety _________________________ 40 

Field of view ______________________ 44, 45 

Filter, optical ________________________ 38 

Footprint __________________________ 41 

Gain control _____________________ __ 38 

Housing ___________________________ 38 

Inner axis ________________________ 36, 37 

Mount structure. See Gimbal, structure, tracker. 

Optics ____________________________ 38 

Outer axis ________________________ 36, 37 

Pointing accuracy _____________________ 38 

Power, peak ________________________ 40 

Precision __________________________ 40 

Pressure transducer ______________ 33-36, 43, 58 

Pulse rate ________________________ 39, 40 

Quadrant photodiode ________________ 38, 40, 43 

Range equation ______________________ 41

Laser tracker-Continued

Range, maximum slant __________________ 40 

Range measurement ______________ 40, 41, 43, 84 

Receiver ___________________________ 38 

Search mode ____________________ 41, 43, 58 

Signal processor ______________________ 43 

Silica gel __________________________ 38 

Sweep angle ____________________ 38, 44, 45 

Temperature transducer ______________ 43, 44, 58 

Time-interval counter ___________________ 43 

Torque motor. See Gimbal.

Track mode __________________ 41, 43, 58, 59 

Transmitter ______________________ 38, 43 

Trigger pulse ________________________ 43 

Wave form _________________________ 40 

Wavelength ________________________ 40 

Window ________________________ 45, 46 

Latitude, astronomical. See Coordinate-axis system, latitude

coordinate.

Latitude coordinate. See Coordinate-axis system. 

Law of Gyroscopics ______________________ 11 

Level. See Locally level platform, Tangent-plane platform. 

Light, velocity of _______________________ 66 

Line printer __________________________ 52 

Local vertical ____________________ 7, 29, 31, 33

See also Gravity-field vector.

Locally level platform ___________________ 9, 24 

Longitude coordinate. See Coordinate-axis system. 

Low-frequency motion ___________________ 7, 8

M

Magnavox receiver ______________________ 86 

Magnetic circuits, gyro ____________________ 12 

Magnetic fields, resolver __________________ 21 

Magnetic flux, resolver ____________________ 21 

Magnetic shielding, resolver ________________ 21 

Magnetic tape ________________________ 54 

Master clock. See Master quartz-crystal oscillator. 

Master operating program. See Computer, software. 

Master quartz-crystal oscillator _____________ 43, 56 

Memory, computer ________________ 8, 49, 50, 51 

Meridian _________________________ 29-31 

Meru. See Milli-earth-rate unit.

Milli-earth-rate unit _____________________ 14

Modes, APT system ___________________ 55-59

Aerial calibrate ___________________ 59, 66, 83

Aerial survey ____________________ 59, 66, 83

Calibrate and aline _______________ 57, 58, 66, 83

Computer off ________________________ 56

En route ______________________ 58, 66, 83

Search ______________________ 58, 59, 66, 83

Standby _______________________ 58, 66, 83

Startup and test ______________ 56, 57, 61, 66, 83

Moisture condensation, optics ________________ 38

Momentum, angular. See Gyro.

N 

National Geodetic Vertical Datum of 1929 ____ 9, 79, 86, 87
102 INSTRUMENT SYSTEM FOR AERIAL SURVEYING

Natural frequency, IMU gimbal _______________ 19 

Navigation _________________________ 7, 58 

NAVSTAR navigation system ________________ 81 

Newton's laws ________________________ 8, 9 

Nonlinear techniques, electronic design __________ 24 

North American Datum of 1927 ____________ 9, 79, 85 

North geographical pole _________________ 29, 30 

Northern hemisphere ____________________ 29 

Null position _____________________ 16, 23, 24

0

Operating altitude, aircraft _________________ 45

Operator terminal. See Computer, digital, flight.

Optical filter _______________________ 38, 70

Optics. See Laser profiler, Laser tracker.

Optimal smoothing. See Data processing.

Orientation in space ________________ 1, 9-11, 29

See also Inertial navigator. 

Orthogonal axes. See Axis, orthogonal. 

Output axis. See Gyro. 

Overshoot, feedback loop __________________ 23

Parallel of latitude ____________________ 29-31

See also Coordinate-axis system, latitude coordinate. 

Parity check, computer ____________________ 50 

Path of motion. See Aircraft, flight path. 

Pendulous integrating gyro accelerometer _______ 17,18

See also Accelerometer.

Pendulum ________________________ 15, 16 

Performance dynamics, APT system __________ 27-35 

Performance specifications, APT system __________ 3 

Photodetector, silicon-avalanche ______________ 38 

Pitch motion ____________________ 7, 19, 24, 78

See also Aircraft motion.

Pitch stabilization loop ____________________ 24 

Plasma display panel _______________ 51, 52, 56, 57 

Platform. See Inertial measurement unit, stable platform. 

Pointing accuracy, tracker __________________ 38 

Polar axis _________________________ 29, 31 

Pole, geographical ____________________ 29, 30 

Position coordinates. ______________ 83-85, 89, 91

See also Coordinate-axis system.

Position error ____________________ 83, 90, 91 

Precession torque. See Gyro. 

Precision, double. See Computer, double precision. 

Precision of measurement. See Laser profiler, Laser tracker. 

Pressure transducer. See Laser tracker. 

Priority level. See Computer, priority level. 

Profiler. See Laser profiler. 

Programs. See Computer, software.

Q

Quadrant photodiode ________________ 38, 40, 43

R

Range equation, laser __________________ 41, 71

Range measurement. See Laser profiler, Laser tracker. 

Real time __________________ 3, 10, 50, 51, 56, 60 

Reference ellipsoid _____________________ 81 

Reflector, corner-cube. ____________ 45, 46, 78, 79, 81

See also Retroreflector.

Renavigating _________________________ 83 

Resolver, accelerometer ___________________ 17 

Resolver encoders ______________________ 24 

Resolver, IMU ______________________ 19-22 

Resolver, tracker ______________________ 38 

Retroreflector ____ 4, 7, 36-38, 41, 43, 45, 46, 58, 59, 77-79,

81, 83, 85, 87-89, 94

Spacing _______________________ 81, 89, 94 

Return-signal pulse ______________________ 38 

Right-hand-thread rule ____________________ 31 

Rockwell Commander aircraft ______________ 6, 77 

Roll gimbal ________________________ 19, 24 

Roll motion ____________________ 7, 19, 24, 78

See also Aircraft, motion.

Roll stabilization loop ____________________ 24 

Rotation, stable platform ______________ 29, 31-33 

Rotor, resolver ________________________ 21

Safety, eye ________________________ 40, 70

Satellite receiver _____________________ 80, 86

Schuler period ________________________ 34

Search. See Modes, APT system.

Second-order surveys _________ _ 81, 86

Self alinement, stable platform ___________ 29, 33, 34

Sensitive axis. See Accelerometer, Gyro.

Sensor, air temperature. See Laser tracker, temperature

transducer. 

Sensor, static-air pressure. See Laser tracker, pressure

transducer.

Servoamplifiers __________________ 7,12, 23-25 

Servomotor. See Gimbal, torque motor. 

Signal generator. See Accelerometer, Gyro. 

Silica-gel _ _ 38 

Silicon-avalanche photodetector ______________ 38 

Software. See Computer, digital, flight. 

Specific force ______________ 9,15,17, 27, 29, 33, 34 

Spin axis. See Gyro. 

Spin velocity, gyro rotor _____________ 11,12

Stable platform. See Inertial measurement unit, stable

platform. 

Stabilization loops _______________ __ 23, 24

State-of-health sensors ____________ __ 24, 58 

State variables, method of _____ _ 54

Stator, resolver __________________ _ 21

Support structure. See Gimbal, structure.

Sweep angle, tracker ____ 38

Tasks. See Computer, software.

Tangent-plane platform _ 9, 27, 29, 33 

Tape, magnetic ___ 54, 83 

Telemetry system _____ 24, 61
INDEX 103

Temperature control _________________ 10, 25-27 

Temperature transducer. See Laser tracker. 

Terminal, operator's. See Computer, digital, flight. 

Terrain profiles __ 1-4, 36, 66, 70, 73, 78, 81, 83-85, 87-89,

91, 92

Thermal environment ________________ 10, 25-27 

Thermal shims, gyro _____________________ 12 

Third-order surveys _____________________ 81 

Tilt, stable platform _______________ 23, 24, 33, 34 

Time characteristic, linear system ______________ 14 

Time-code signal _____________________ 59, 61

See also Master quartz-crystal oscillator. 

Time-date generator _____________________ 59 

Time-interval counter ____________________ 43 

Titanium alloy, housing, tracker ______________ 38 

Torque. See Accelerometer, Gyro. 

Torque motor. See Gimbal. 

Tracker. See Laser tracker.

Transimpedance preamplifier __________ 38 

Triangulation stations _____ __ 80, 81 

TV camera. See Color TV camera. 

Twin Otter aircraft. See Aircraft.

U

UNIBUS. See Computer, digital, flight.

Update, inertial navigator ________ 4, 7, 35, 36, 83, 87-91

Variable-speed fan ______________________ 27

Velocity measurement. See Accelerometer.

Vertical acceleration _____________________ 35

.S. G.P.O. 1987-181-408:4000?

Vertical accuracy _________________ 1, 81, 89-92 

Vertical control ____________________ 1, 77-79 

Vertical coordinate. See Coordinate-axis system, elevation

coordinate.

Vertical velocity _______________________ 34 

Video imagery ________________ 78, 83, 84, 92, 93 

Video recorder ____________________ 59, 73, 75 

Video subsystem. See Color TV camera, Video recorder. 

Viscous damper, gyro __________________ 10,11 

Viscous fluid, gyro ____________________ 12, 14 

Viscous shear, gyro ______________________ 14

W

Wander, gyro ________________________ 14

See also Gyro, precession.

Wheel inertia, gyro ______________________ 11 

Window __________________________ 45, 46

X

x axis. See Cartesian coordinates, Coordinate-axis system.

Yaw motion ______________________ 7, 19, 24

See also Aircraft, motion. 

y axis. See Cartesian coordinates, Coordinate-axis system.

z axis. See Cartesian coordinates, Coordinate-axis system.