The presence of failures in navigation sensors can cause the determination of an erroneous aircraft state estimate, which includes position, attitude, and their derivatives. Aircraft flight control systems rely on sensor inputs to determine the aircraft state. In the case of integrated Inertial/NAVSTAR Global Positioning System (GPS), sensor failures could occur in the on-board inertial sensors or in the GPS measurements. The synergistic use of both GPS and the Inertial Navigation System (INS) allows for highly reliable fault detection and isolation of sensor failures. Integrated Inertial/GPS is a promising technology for the High Speed Civil Transport (HSCT) and the return and landing of a manned space vehicle

Kline, Paul

Vangraas, Frank

English

NASA Technical Reports Server

INTEGRATED INERTIAL/GPS
Paul Kline and Frank van Graas
Ohio University
Athens, Ohio
SUMMARY
The presence of failures in navigation sensors can cause the determination of an erroneous
aircraft state estimate, which includes position, attitude, and their derivatives. Aircraft flight
control systems (AFCS) rely on sensor inputs to determine the aircraft state. In the case of
integrated Inertial / NAVSTAR Global Positioning System (GPS), sensor failures could occur
in the on-board inertial sensors or in the GPS measurements. The synergistic use of both GPS
and INS allows for highly reliable fault detection and isolation of sensor failures. Integrated
Inertial/GPS is a promising technology for the High Speed Civil Transport (HSCT) and the
return and landing of a manned space vehicle.
BACKGROUND
The NAVSTAR Global Positioning System is expected to be operational by the year 1993.
Presently, ten experimental satellites (Block I) and seven operational satellites (Block II) have
been successfully launched and are undergoing tests. Navigation with the GPS is based on
pseudorange and integrated Doppler shift measurements with respect to GPS satellites. In
combination with ephemeris information, transmitted by the satellites, a receiver can determine
its position and velocity with respect to the satellites with a positional accuracy of 16 meters
(Precise Positioning Service) and with a velocity accuracy of better than 0.1 meter per second.
Differential GPS can result in navigational accuracies of better than 3 meters for
pseudoranging and well within 0.2 meter using integrated carrier phase measurements. The
GPS satellite signals can also be used to determine aircraft attitude by making interferometric
measurements with a minimum of three separate antennas (ref. 1). GPS interferometric
attitude determination is expected to be within 0.1 degree for pitch, roll, and yaw angles.
Because of the capabilities described above, the GPS can play an important role in the
reliable state estimation of aircraft. Specifically, an integrated Inertial/GPS system could
provide continuous, high-accuracy attitude information, three-dimensional position and velocity,
and time. The system could also provide a high level of fault tolerance, since GPS and inertial
measurements provide similar information.
The overall objective of an integrated Inertial/GPS system is to provide reliable aircraft state
estimates in the presence of possible failures in the GPS measurements and the on-board
inertial sensors. This involves the design of fault tolerant algorithms, implementation and flight
demonstration of the integrated Inertial/GPS system. Previous research at Langley Research
Center demonstrated the capability to provide dual fail-operational performance for a skewed
array of inertial sensors (ref. 2). The redundant strapdown inertial measurement unit
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https://ntrs.nasa.gov/search.jsp?R=19910009719 2020-03-24T06:46:36+00:00Z
(RSDIMU) used for this project will also be used fl_rthe lalx_ratory and flight demonstration
of lnertial/GPS techniques.
OVERVIEW OF RESEARCH
Current efforts are focused on the determination of the effective GPS data rate which
involves GPS receiver tracking loop design and simulation, and on fault detection and isolation
of GPS measurement errors. In the area of receiver autonomous fault detection, a novel
signal processing technique has been applied to improve the effectiveness of fault detection
algorithms in the presence of near collinearity (ref. 3).
The GPS receiver simulation program as presented in reference 1 is being extended to
address the effect of tracking loop bandwidths of the code and carrier tracking loops on the
effective GPS data rate. Figure 1 shows the block diagram of an integrated lnertial/GPS in
combination with the AFCS. Both GPS and inertial data are combined by the estimator,
which in turn provides the aircraft state estimates to the flight control system. The GPS data
rate is required to facilitate the integration of GPS data into the aircraft flight control system,
and to determine fault detection capabilities based on GPS attitude information.
Also continuing is the implementation of tlre integrated lnertial/GPS test bed. Tire
RSDIMU system is currently available for data collection in the laboratory environment. The
inertial system will be combined with a 4-channel GPS receiver. At the same time, provisions
are being made to install the integrated system in Ohio University's Douglas DC-3 research
aircraft, N7AP.
REDUNDANT STRAPDOWN INERTIAL MEASUREMENT UNIT
The inertial sensor used for this project is a Redundant Strapdown Inertial Measurement
Unit (RSDIMU). A block diagram of the RSDIMU is shown in figure 2 (based on ref. 4).
The inertial sensors consist of four two-degree-of-freedom accelerometers and gyroscopes
arranged in two semi-octahedral arrays. These redundant arrays are interfaced to electronics
boxes which contain analog-to-digital converters. The sensor data are processed by two flight
computers. The magnetic tape drive is used for in-flightprogram loading. Also used during
system testing is a printer for real time data output and a control panel to operate the system
and to inject failures into the measurement data. Power to the inertial sensors is provided
through a power back-up unit, which allows the sensors to operate for about 20 minutes in the
absence of aircraft power.
The performance of the inertial sensors would support navigation accuracies on the orcter of
2 nautical miles per hour. However, the inertial system is only required to provide slmrt-term
navigation data. The GPS will provide the long-term navigation stability for this system.
RSDIMU NAVIGATION MODE PROGRAM FLOW
The RSDIMU has several modes of operation, including coarse and fine alignment, and
navigation. Figure 3 shows the flow diagram for the navigation mode. Each navigation cycle
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is initiated by a sensor interrupt. The raw sensor data are then scaled and converted to floating
point representation. Next, fault detection and isolation routines are carried out, and some
data are fed back to the IMU sensors for corrections. Finally, attitude informaticm is output to
the flight control system and the navigation solution is updated. The subsequent sensor
interrupt initiates the next run through the navigation update cycle. The system operates at an
update rate of 64 updates per second.
GPS RECEIVER
A four-channel GPS receiver employing continuous carrier tracking is used for this project.
Figure 4 shows a simplified block diagram of the GPS receiver tracking loops. The bandwidth
of the code tracking loop is as small as 0.01 Hertz when the receiver is in continuous tracking.
The carrier tracking loop bandwidth is 32 Hertz and is used to update the code tracking loop
every millisecond to correct for dynamics which cannot be tracked by the code loop. The GPS
carrier wavelength at 1575.42 MHz is approximately 19 cm, which provides a tracking
resolution of better than 1 cm. In addition, the carrier tracking loop is a second order system
which has less than 20 degrees of phase lag at line-of-sightaccelerations of up to 40 m/s',
which corresponds to a tracking error of less than 1 cm. The continuous carrier tracking is
essentially a dead-reckoning input, similar to Doppler radar.
APPLICATIONS
The integration of GPS and INS has been targeted for several possible applications. These
applications actually define the overall system requirements. One application is the High
Speed Civil Transport (HSCT). GPS/INS can be used for heading and attitude determination.
The return of a Space Shuttle from orbit or a Manned Return Vehicle (MRV) from a space
station is another application. The inertial measurement would provide the short-term stability
required by the flight control system. Differential GPS would be used in the terminal area to
greatly improve landing accuracy and to enable the use of candidate landing sites throughout
the world without the need for a dedicated landing system installation.
CONCLUSIONS
The synergistic integration of inertial measurements and the GPS operated stand-alone or in
the differential mode has many promising applications in the areas of reliable aircraft state
estimation and for use in terminal and landing operations. The RSDIMU test bed is
operational in the laboratory environment, and is being integrated with a GPS receiver to
implement and evaluate the above applications and to develop fault detection and isolation
algorithms.
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REFERENCES
1. Mclntyre, D. S.: GPS Effective Data Rate Characterization with Application to Integrated
GPSflNS Attitude and Heading Determination. Master'sThesis, Ohio University, Department
of Electrical and Computer Engineering, Athens, OH, November 1989.
2. Morrell, F. R., Bailey, M. L., and Motyka, P. R." Flight Demonstration of Redundancy
Management Algorithms for a Skewed Array of Inertial Sensors. Proceedings of the
AIAA/AHS/ASEE Aircraft Design and Operations Meeting, AIAA Paper 88-4434,Atlanta,
GA, September 7-9, 1988.
3. Kelly, R. J., van Graas, F., and Kuhl, M. K.: Improved Effectiveness of GPS RAIM
through Ridge Regression Signal Processing. Proceedings of the 2nd International Satellite
Division Meeting of the ION, Colorado Springs, CO, September 27-29, 1989.
4. Bryant, W. H. and Morrell, F. R.: Description of a Fail-Operational Redundant Strapdown
Inertial Measurement Unit for AvicmicsResearch. Proceedings of the 4th ROLM Computer
Users Group Meeting, San Diego, CA, February 1981.
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GPSSIGNA_ "1
l_ON'r- ._D
i GPS
RECEIVER
I RE_ ICONTROLS -_ BANDWIDTHADJUSTMENT
GPS DATA AT TIE
EFFEC'rlVE DATA
RATE, Roq
CONTROLS
t 'V----1 _r_AS_ p----- DYNAMICS
YAW VELOCITY DATA
cOWrROL
SYSTEM
ffcs)
Figure 1. Integrated Inertial/GPS in combination with the Aircraft Flight Control System.
TAPE
_ UNIT
I / NAVIGATION
IMU I I FLIGHTl.,.4--_Ib.-CONTROL
A 1-,_-----I_ COMPUTER [ PANEL
I FIJGIITI P_ER I I TERMINALI I COMPUTER
MEASUREMENTIlMuIUNIT CONDITIONER
COMPUTER
PROCESSOR
B
B COMPUTER _. t t . COMPUTER
PROCESSOR B CONTROL PANEL
1/O I/O B
Figure 2. Block diagram of the Redundant Strapdown Inertial Measurement Unit (RSDIMU)(based on ref. 4).
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l
WAIT FORSENSOR INTE RUPT
I co_TS_SO_IDATATO FLOATINGPOINTAND SCALE
1
FAULTDETEUFION [
AND ISOLATION
1
1
DATA OUTPUT
1
NAVIGATION [
ALGORFI'HM
I
Figure 3. RSDIMU navigation mode flow diagram.
0,01 |1: BANDWlDTI! 0.5second
CODE ] 3-5m NOISE PSI_III'_'_RANGF
.__ J_
TRACKING I---
LOOP
CO RP,.ELATOR ( )
TRACKING _"
LOOP (DELTA
PSEUDORANGE)
•41 0.1 mNOISE
CORRECTIONSTO CODEPOSITION
Figure 4. GPS receiver simplified block diagram of the code and carrier tracking loops.
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