interval. For a low performance system the noise level 
will be above the one second value. 
  
  
  
  
  
Error in System Accuracy Class 
nav. grade low accuracy 
Attitude pitch & roll; ^ azimuth pitch & roll; azimuth 
1h 10"-30"  60"-180" 0.5 - 1.0 49.- 3? 
1 min 5" - 10" 15" - 20" 01-03 | 02-05" 
1s 3-5 3" - 20" 0.01-0.02 .02°-0.05° 
Velocity 
1h 0.5 - 1.0 m/s 200 - 300 /s 
1 min 0.03 - 0.10 m/s 1-2m/s 
1s 0.001 - 0.003 m/s 0.002 - 0.005 m/s 
Position 
1h 500 - 1000 m 200 - 300 km 
1 min 0.3-1.0m 30-50m 
1s 0.02 - 0.05 m 0.3-0.5m 
  
  
  
  
  
Tab. 2: Current INS Performance. 
4. ACCURACY ACHIEVABLE BY AN INTEGRATED 
INS/GPS 
As is obvious from Tables 1 and 2, the stand-alone 
accuracy of each system will not give the highest 
possible accuracy. INS will have superior orientation 
accuracy, GPS superior position accuracy. Thus, an 
integration of the two systems will result in an optimal 
solution which will also provide much needed 
redundancy. 
GPS positioning using differential carrier phase is 
superior in accuracy as long as no cycle slips occur. 
GPS relative positions are, therefore, ideally suited as 
INS updates and resolve the problem of systematic error 
growth in the IMU trajectory. On the other hand, the IMU- 
derived attitude is usually superior to that obtained from 
a GPS multi-antenna system. In addition, IMU-derived 
position differences are very accurate in the short term 
and can thus be used to detect and eliminate cycle slips 
and to bridge loss of lock periods. Because of the high 
data rate, they provide a much smoother interpolation 
than GPS. Integration of the two data streams via a 
Kalman filter thus provides results which are superior in 
accuracy, reliability, and homogeneity. 
The following figures will illustrate the orientation 
accuracy currently achievable with an integrated 
INS/GPS. Position accuracy is not discussed in the same 
detail because it is largely dependent on GPS accuracy 
for which Table 1 and the references given there can be 
consulted. 
Fig. 2a shows the attitude output of a navigation-grade 
INS in static mode over a period of about 25 minutes. 
The systematic error is smooth and reaches a minimum 
after about 21 minutes. It is clearly part of the Schuler- 
type oscillation. The noise about this trend is very small 
and shows white noise characteristics. After eliminating 
the trend, the noise pattern in Fig. 2b results. It shows 
the system attitude noise which is at the level of 3 
arcseconds (RMS) and represents the attitude accuracy 
achievable under ideal conditions. Fig. 2c shows the 
noise of the same system mounted in an aircraft with the 
engines switched on. As before, the trend has been 
eliminated. In this case, the low noise level of the lab test 
cannot be maintained. The noise is now between 15 and 
  
  
  
  
0.093 
0.02} 4 
E 0.01} - 
"B 
Or 4 
-0.02 ic = 
time (=) 
Figure 2a: Total roll error, static case, lab 
  
  
  
  
time (©) 
Figure 2b: Roll noise, static case, lab (G =3") 
  
0.02 T T T T 7d, T T T 
roll error (deg) 
     
  
E 
Jj 
   
  
  
? 1 1 1 1 1 L 1 1 
eO 600 700 800 S900 1000 1100 1200 1300 1400 
time (sec) 
Figure 2c: Roll error, static case, aircraft, engine 
switched on (O - 18") 
0.02 T T T T T T T T 
  
e 
© 
Un 
T 
1 
roll error (deq) 
C 
  
  
  
  
0020 600 700 E00 S00 1000 1100 1200 1300 1400 
time (sec) 
Figure 2d: Roll error, static case, aircraft, engine on, 
vibration filter (11") 
70 
International Archives of Photogrammetry and Remote Sensing. Vol. XXXI, Part B6. Vienna 1996 
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