re) ar © ron) Py ”) o o © " pm wu) fn)
BALLISTIC  MISSILE  GUIDANCE  AND 
CONTROL  -  USSR  AND  PRC 
DST-100QS-293-76 
DIA  TASK   NO.   PT- 1000-01-75 
DATE  OF  PUBLICATION 
12  MAKCH  1976 
Information  Cutoff  Date 
9  June  1975 
This    study    supersedes    ST-CS-0S-293-75,    "Ballistic    and    Defensi, 
Missile   Guidance  and  Control—  ECC  (U),"  dated  14  March   1975. 
*  ^\ 
(Reverse  Blank)
DST-1000S-293-76 
12  March  1976 
PREFACE 
;        The  following  definitions  apply  to  terms  used  in  this  study: 
1.  Estimate  -  Judgment  regarding  the  present  or  a  historical  period  about 
which  there  is  substantial  validated  information;  extended  to  include  the  next  2  years. 
2.  Projection  -  Judgment  regarding  the  time  frame  from  2  years  hence  with 
decreasing  confidence  as  the  time  span  increases. 
v 
3.  Almost  Certain  -  Indicates  approximately  a  90:10  chance  that  an  event 
will  occur.  Related  terms:  will,  shall,  is  expected,  is  anticipated. 
4.  Probable  -  Indicates  approximately  a  75:25  chance  that  an  event  will 
occur.  Related  term:  likely. 
5.  50:50  -  Even  chance  an  event  will  occur. 
6.  Unlikely  -  Indicates  approximately  a  25:75  chance  or  less;  i.e.,  only  one 
chance  in  four  it  will  happen.  Related  term:  improbable. 
7.  Better  Than  Even  Chance  -  Greater  than  a  50:50  chance  but  less  than 
approximately  a  75:25  chance. 
8.  Less  Than  Even  Chance  -  Less  than  a  50:50  chance  but  greater  than 
approximately  a  25:75  chance. 
This  revision  contains  significantly  new  and  revised  information  which  affects 
previously  published  assessments.  A  delta  (A)  preceding  a  paragraph,  heading,  or 
caption  indicates  that  the  discussion,  table,  or  figure  is  reporting  significantly  new  or 
revised  information. 
in 
(Reverse  Blank)
DST-1000S-293-76 
12  March   1976 
METRIC  CONVERSION  TABLE 
|  Parameters  and  measurements  in  this  document  are  reported  in  metric  units,  to  bring 
Department  of  Defense  intelligence  reporting  into  consonance  with  internationally  accepted  standards 
for  measurement.  For  the  benefit  of  users  who  require  parametric  values  expressed  in  customary  units, 
the  accompanying  table  lists  each  metric  unit  applicable  to  this  document,  along  with  its 
corresponding  value  in  customary  units. 
[  The  metric  system  used  is  the  International  System  of  Units  (SI),  along  with  certain  non-SI 
metric  units  approved  by  the  International  Committee  on  Weights  and  Measures  (ICWM)  for  use  with 
the  SI  system.  Certain  non-metric,  international  customary  units  such  as  degree  of  arc,  and  the 
customary  units  of  time  also  have  been  approved  by  ICWM  for  continued  use  with  the  SI  system  and 
will  continue  to  be  employed  where  appropriate.  US  customary  units  also  will  continue  to  be  employed 
for  items  which  are  defined  or  named  in  terms  of  customary  units,  such  as  1-megaton  warhead,  .'38- 
caliber  pistol,  and  two-by-four.  An  asterisk  (*)  in  the  Customary  Unit's  column  indicates  that  the 
customary  unit  will  continue  to  be  used  for  the  indicated  parameter. 
PARAMETER 
VALUE  IN                                                       METRIC 
METRIC               METRIC  UNIT             SYMBOL 
ACCELERATION,  SPEED,  VELOCITY 
VALUE  IN 
CUSTOMARY 
CUSTOMARY 
UNIT 
Acceleration 
Velocity 
meter/second*                     m/s1 
meter/second                      m/s 
3.280840 
3.280840 
ft/sec* 
ft/sec 
RANGE,  ALTITUDE,  DIMENSION,  LENGTH 
Range/Distance 
Altitude  dimension 
Dimension 
Thickness,  wavelength 
kilometer                           km 
meter                                  m 
millimeter                          mm 
micrometer                          fim 
MASS  (WEIGHT) 
0.5399568 
3.280840 
0.03937008 
0.03937008 
NM 
foot 
inch 
mil 
Gross  weight,  payload 
Mass  (bulk) 
I 
kilogram                            kg 
metric  ton  (tonne)             t 
2.204623 
1.102311 
pound  (m) 
ton 
ENERGY  (J- W  •  »-N  •  m),  TORQUE  (N  •  m) 
Torque 
newton-meter                     N  •  m 
0.7375621 
lbf.ft 
PRESSURE,  STRESS,  STRENGTH  (Pa-N/m») 
Pressure  (gage) 
Pressure  (absolute) 
kilopascal                            JtPa  (gage) 
kilopascal                            kPa  (abs) 
ANGLE 
0.14S0377 
0.1450377 
psig 
psia 
Plane  angle 
Angular  velocity 
I 
radian                                  rad 
radian/second                     rad/s 
57.29579 
9.549279 
•degree 
*rprn 
{Reverse  Blank)
DST-lOOOS-293-76 
12  March  1976 
TABLE  OF  CONTENTS 
Page  No. 
Preface 
  iii 
Metric  Conversion  Table    v 
Summary    xiii 
Section    I     Soviet  Surface-to-Surface  Ballistic  Missile  Guidance  and  Control           1 
1.  Introduction    1 
2.  ICBM  Guidance  and  Control  •   „ 
a.  SS-7    2 
b.  SS-8              
c.  SS-9              
d.  SS-11           
e.  SS-13            
—  f.      SS-X-15    38 
•  —  g.     SS-X-16         
h.     SS-17            
i.      SS-18          
 : 
j.      SS-19         
3.  SRBM,  MRBM,  and  IRBM  Guidance  and  Control';       
a.  SCUD  A  and  SCUD  B          
b.  SS-4          
 : 
c.  SS-5  ,         
d.  SS-12    72 
—  e.      SS-14           
4.  Naval  Ballistic  Missile  Guidance  and  Control        
a.  SS-N-4           
 '.    75 
b.  SS-N-5    77 
c.  SS-N-6 
d.  SS-N-8                 
e.  SS-NX-13             
5.  Trends         .„ 
 „ 
a.  Inertial  Component  Capabilities  and  Trends    87 
b.  Soviet  Ballistic  Missile  Guidance  Technology  Trends      "  
c.  SRBM  Guidance  and  Control  Trends    94 
d.  MRBM/IRBM  Guidance  and  Control  Trends            95 
e.  ICBM  Guidance  and  Control  Trends        
f.  Postulated  Reentry  Guidance        
vu
DST-1OO0S-293-76 
12  March  1976 
TABLE  OF  CONTENTS  (Cont)  • 
Pago  No. 
Section  II     People's  Republic  of  China  Surface-to-Surface  Ballistic  Misaile 
Guidance  and  Control  j       
  105 
1.     Technology  Background  /      
       105 
a.  Computer  Developments  '         105 
b.  Missile  Guidance  Radar  Development  ,'       
       105 
c.  Automatic  Control  Components  |     ' 
  106 
d.  Miniaturization  and  Semiconductors  Development  f    106 
e.  PRC  Commerical  Trade  .       
  106 
2.  System  Control  Theory  '     
  106 
3.  Guidance  Assessments  {    107 
4.  Component  Assessments    „ 
  107 
5.  Error  Analysis  of  the  CSS-2  IRBM  System  /       
 Z!1"Z!11ZZZ!Z1Z"  107 
6.  Guidance  Accuracies  j       ...-. 
  HO 
Appendix       I     USSR  and  PRC  Missile  Guidance  Systems  i      . 
 '.    H3 
Appendix     II     Principal  Error  Axis  Concept/      
 __  H7 
Appendix    III     Cutoff  Law  Mechanizationa.'      
  12i 
Appendix    IV    Guidance  Errors  j       
 ' 
  123 
Appendix    VI    Soviet  Naviational  Aids  |     .' 
  127 
Appendix  VII     Pendulous  Proportional  Gyro  Accelerometer  |       
  131 
Bibliography 
  133 
LIST  OF  ILLUSTRATIONS 
Figure     1     Missile   Guidance   Functional   Diagram  f       
Figure    5    Miss  va  Impact  Range — Guidance  Contribution                   17 
Figure     6     CEP  vs  Impact  Range — Guidance  Contribution            18 
vm
DST-1000S-293-76 
12  March  1976 
LIST  OF  ILLUSTRATIONS  (Cont) 
Page  No. 
Figure     7     Miss  vs  Impact  Range — Guidance  Contribution                   20 
Figure     8     CEP  vs  Impact  Range — Guidance  Contribution    "             21 
Figure  12     Variable  Gain  Actuation  System       ,           28 
Figure  15  Modified  PEA  Guidance  Mechanization  {    34 
Figure  16  Pitch  Steering  Loop         
 -    35 
Figure  17  Nominal  Pitch  Command  j    35 
Figure  18  Stable  Platform  After  Pitchover         
s 
Figure  20    Accelerometer  Orientations  {      
Figure  21     Guidance  and  Control  Functional  Block  Diagram                   41 
Figure  27     Guidance  Coordinate  System 
V   / 
IX
\ 
DST-lOOOS-293-76 
12  March  1976 
LIST  OF  ILLUSTRATIONS  (Cont) 
Pago  No. 
Figure  31     Guidance  System  Mechanization  \       
Figure  34    SCUD  Guidance  and  Control  System  '       
V    f 
Figure  36    Terrain  Navigator  Used  with  SCUD  System  v        
Figure  42     Estimated  Trend  in  Soviet  Stabilization  Gyro  Technology           93 
Figure  43    Accelerometer  Quality  Trend  Estimates  \      
Figure  44    MRBM/IRBM  Accuracy  Trend  Estimate           96 
Figure  45    ICBM  Inertial  Guidance-Only  Accuracy  Trend         
a 
u 
-  i\ 
□i 
u
/" 
DST-1000S-293-76 
12  March  1976 
LIST  ILLUSTRATIONS  (Cont) 
Page  No. 
Figure  50    Downrange  Miss  vs  ATbo— 2,950-krn  Trajectory           109 
Figure  51     Cutoff  Mechanization — Time  Compensation  System           110 
Figure  52    Trajectory  Coordination  System    117 
Figure  53     Principal  Axis  Coordinate  System  
  118 
Figure  54     Soviet  Directional  Gyroscopes          
  125 
Figure  56     Diagram  of  Pendulous  Proportional  Gyro  Accelerometer  (U)    131 
LIST  OF  TABLES 
Table  I    Accelerometer  Orientations  and  Pitch  Attitude  at  Burnout 
(Degrees)    "*   ...-.    4 
Table            II    First  Stage  Pitch  Program         
Table           III     Accelerometer  Orientations    13 
Table  XI    Relative  Capability  of  Available  Navigation  Aids 
Table         XII    Soviet  Inertial  Component  Technology  Capabilities                  89 
XI
DST-1000S-293-76 
12  March  1976 
LIST  OF  TABLES  (Cont) 
Page  No. 
Table  XVI    Three-Sigma  Parameter  Uncertainties  (U) 
  108 
Table  XVII    Guidance  Errors  For  System  Without  Time  Compensation  (U)    108 
Table  XVIII    Time  Compensation  Parameters  (U)    110 
Table  XIX    Guidance  Errors  for  System  With  Time  Compensation  (U)    Ill 
Table  XX     Guidance  System  Accuracy  (U) 
  Ill 
Table  XXI     Offensive  Ballistic  Missile  Guidance  Systems  (U) 
  115 
Zll
DST-1 0003-293-76 
12  March  1976 
SECTION  I 
SOVIET  SURFACE-TO-SURFACE  BALLISTIC  MISSILE 
GUIDANCE  AND  CONTROL  (U) 
1.     Introduction  \ 
1  ! 
i  I 
1   ,  I 
i        A  method  of  improving  the  accuracy  of  an  inertial  guidance  system,  especially  mobile 
launchers,  is  to  use  a  star  tracker.  A  star  tracker  can  correct  for  both  position  and  azimuth  errors.  It 
can  greatly  reduce  the  need  for  frequent  updatings  of  the  navigation  system  on  the  mobile  launcher 
The  amount  of  improvement  depends  on  the  quality  of  Ihe  inertial  components  ^sed  in  the  inertial 
guidance  system  on  the  missile  and  in  the  navigational  system  used  m  the  mobile  launcher.
l//  -  Cj  u  +  c2  u  +  c3«  ♦  c4 
DST-1O00S-293-76 
12  March   1976 
NORMAL 
ACCELERO- 
METER 
FTD  A70-2102 
Fig.  16     Pitch  Steering  Loop 
IHH UST    ANGLt 
lutui 
«. 
{ 
— 
\ 
  S 
v-* 
. 
«^- 
20  40  60  80  100  120  140  160 
TIME  ISECJ 
FTD  A70-2103 
Fig.  17     Nominal  Pitch  Command
DST-1000S-293-76 
12  March   197^ 
\ 
TILT  MOTOR 
TILT  TABLE 
PLATFORM 
FID  A72-1249 
Fig.  18     Stable  Platform  After  Pltchover/
DST-1000S-293-76 
12  March  1976 
CROSSRANGE 
HORIZONTAL 
O-    UPPER  INSTRUMENTS  ORIENTATION 
$-    LOWER  INSTRUMENTS  ORIENTATION 
FTD  A75-1883 
Fig.  20    Aceelerometer  Orientations
DST-1000S-293-76 
12  March   1976 
A  possible  guidance  system  limitation  does  exist  for  any  future  MIRV  use;  i.e., 
use  against  targets  hundreds  of  miles  apart.  The  guidance  concept  now  employed  may  not  easily  adapt 
to  MIRV  use.  The  introduction  of  a  high  capacity  digital  computer  and  a  change  in  guidance 
philosophy  will  likely  have  to  take  place  to  obtain  the  flexibility  and  onboard  navigation  capability 
required  for  advanced  multiple  RV  delivery  systems.
DST-1000S-293-76 
12  March  1976 
Gyro  error  sources  (Table  XII)  can  be  separated  into  two  main  categories:  (1) 
systematic  torques  and  (2)  random  torques. 
Systematic  torques  are  defined  as  torques  which  can  be  measured  and  whose 
characteristics  can  be  correlated  with  some  parameter.  It  is  theoretically  possible  to  compensate  either 
directly  or  indirectly  for  the  errors  produced  by  the  systematic  torques.  The  drifts  due  to  systematic 
torques  are  basically  of  two  types:  (1)  acceleration-sensitive  drifts  generally  caused  by  mass 
unbalance,  and  (2)  acceleration-squared -sensitive  drifts  generally  caused  by  structural  compliance- 
anisoelasticity. 
Random  or  unpredicted  variations  will  introduce  uncompensated  errors  which 
must  be  charged  to  the  gyro.  These  errors  are  defined  as  "constraint"  torques  and  are  the  most 
significant  errors.  "Constraint"  errors  are  the  limiting  accuracy  factor  affecting  gyro  performance.
DST-1000S-293-76 
12  March   1976 
A  TABLE  XII 
SOVIET  INERTIAL  COMPONENT  TECHNOLOGY  CAPABILITIES 
The  literature  indicates  that  the  Soviets  have  a  good  mastery  of  the  theory  of  gas 
bearings,  and  the  techniques  they  are  reportedly  investigating  closely  parallel  European  and  US 
designs.  There  are  many  Soviet  publications  on  gas  bearings  dating  back  as  far  as  the  early  1950's.
DST-1000S-293-76 
12   March   1976 
t 
I 
Successful  developments  of  these  instruments  could  lead  to  their  use  in  ballistic 
missile  guidance  systems  where  high  accuracy,  low  cost,  ease  of  maintenance,  and  reliability  are 
extremely  important. 
The  technical  literature  on  the  ESG  has  indicated  the  Soviets  have  a  good 
understanding  of  its  potential  capabilities.  Since  the  late  1960's,  the  Soviets  have  progressed  rapidly 
from  treating  the  basic  operating  principles  of  suspension  to  analyzing  the  more  complex  problems 
such  as  resistance  to  vibration  and  acceleration,  manufacturing  tolerances,  deformations  under 
rotation,  residual  friction  effects  upon  duration  of  spin,  pickoff,  and  control  torquing;  they  appear  to 
be  well  along  in  the  development.  It  is  apparent  that  they  recognize  that  a  drift  rate  of  0.0001  degree 
per  hour  can  be  achieved  with  the  ESG.
DST-1000S-293-76 
12  March   1976 
A  TABLE  XVI 
THREE-SIGMA  PARAMETER  UNCERTAINTIES 
PARAMETER 
Specific  Impulse 
Initial  Weight 
Mass  Flow  Rate 
Thrust  Misalignment 
In-Plane 
Cross- Plane 
3-SIGMA  UNCERTAINTY 
2.0% 
3.0% 
2.5% 
0.1  deg 
0.1  deg 
A  TABLE  XVII 
GUIDANCE  ERRORS  FOR  SYSTEM  WITH  TIME  COMPENSATION 
IMPACT  RANGE 
DR  MISS 
CR  MISS 
CEP 
(km) 
(km) 
(km) 
(km) 
480 
1.41 
0.30 
1.00 
960 
2.78 
0.67 
2.00 
1,870 
4.72 
1.46 
3.46 
2,950 
6.22 
2.59 
4.70 
108
DST-1000S-293-76 
12  March  1976 
DOWNRANGE  MISS  (km) 
x 
X 
"X 
*       'I 
8   - 
- 
4   - 
'2-- 
  i*   
-2.-    *,, 
■4- 
-6- 
-8 
•10 
-12 
-14  ■ 
.IS  - 
-17 
*  Tbo  (sec) 
* 
as 
FTD  A7S-18S4 
A  Fig.  50    Downrange  Mist  vs  AT*  •  2,950-km  Trajectory  (U) 
109
DST-lGOOS-293-76 
12  Miirch   1976 
A  TABLE  XIX 
GUIDANCE  ERRORS  FOR  SYSTEM  WITH  TIME  COMPENSATION  (U) 
IMPACT  RANGE 
DR  MISS 
CR  MISS 
CEP 
(km) 
(km) 
(Itm) 
(km) 
480 
0.22 
0.44 
0.39 
960 
0.33 
0.93 
0.72 
1.870 
0.41 
1.98 
1.41 
2.950 
0.63 
3.28 
2.30 
l     * 
111 
(Reverse  Blank)
DST-1000S-293-76 
12  March  1976 
APPENDIX  II 
PRINCIPAL  ERROR  AXIS  CONCEPT 
The  principal  error  axis  (PEA)  concept  is  presented  in  this  Appendix.  Figure  52  portrays  the 
launch  centered  inertial  (LCI)  computational  coordinate  system  used  where  Y  is  along  the  geodetic 
vertical,  X  is  directed_downrange,  and  the  Z  axis  forms  a  right-hand  coordinate  system.  The  vehicle 
thrust  axis  is  in  the  X,  Y  plane.  The  range  from  launch  to  impact  is  a  function  of  the  position  and 
velocity  vectors  at  thrust  termination.  For  the  purposes  of  this  discussion,  it  will  be  assumed  that  the 
earth  is  nonrotating;  therefore,  a  flight  time  dependency  is  not  explicitly  included.  However,  in  the 
results  to  follow,  flight  time  is  implicitly  included.  Also,  the  crossrange  effects  will  not  be  explicitly 
handled  in  the  development  of  the  range  control. 
FLIGHT     TR4 
J£CTi 
ORy- 
INJECTION 
TO  TARGET 
FTD  A7L2515 
UNCLASSIFIED 
Fig.  52    Trajectory  Coordination  System  (U) 
(U)  Taking  the  injection  condition   of  a  reference  or  nominal  trajectory,  a  Taylor   Series 
Expansion  of  the  range  equation  can  be  developed  as  follows: 
AR 
M  AX  +  -^  AY  +  ^5  AX  +  -^  AY  + 
ax  3Y 
ax 
9Y 
(1) 
where 
AX  =  X  -  Xo  =  difference  in  X  coordinate  from  the  reference  (nominal)  condition  X» 
Other  A  terms  are  defined  analogously. 
Only  first  order  terms  are  retained  The  principal  position  axis  coordinate  system  is  developed  by 
defining  H  as  a  rotation  about  the  Z  axis  of  the  LCI  coordinated  system  as  portrayed  in  Figure    53. 
117
BST-1000S-293-76 
12  March  1976 
FTD  A71-25U 
Fig.  53     Principal  Axis  Coordinate  System 
The  transformation  of  LCI  errors  into  this  coordinate  system  is 
The  inverse  transformation  is 
AYP 
AX 
cos  $P  sin  Op 
-sin  Op  cos  Sp 
cos  Op  —sin  0P" 
AY  sin  0p      cos  0P 
AX 
AY 
AXP 
AYP 
(2) 
Substituting  equation  (2)  into  position  elements  of  equation  (1)  results  in 
AR  «=  (|«    cos  Op  +    |*  sin  Op)  AXp  -    (|«    sin  Op  -    |*  cos  Op)  AYP  (3) 
Defining  the  XP  axis  as  the  principal  position  axiB,  the  angle  6  P  of  this  axis  is  determined  by  setting  the 
component  of  AXP  to  zero.  Thus 
t,.,    .    3R/3Y 
ten  *'      aWax 
and 
sin  Op  — 
3R/3Y 
;  ,    C08    Op 
3R/3X 
V(tfrnw   "  Jaw +  (!£)■ 
Using  these  in  equation  (3)  results  in 
(4) 
AR  -  -|£  AXp 
118
DST-1000S-293-76 
12  March   1976 
where 
1JT  "    V(3X)     +    ("2)T)    obta,n  ''-v  US!ng  (4)  in  <3>- 
With  the  same  approach,  the  principal  velocity  axis  is  developed. 
(U)  Summarizing  the  above,  the  necessary  formulas  are: 
0,  -  tan-    3R/3Y.  *v  -  tan-  .3*3* 
3R/3X  DR/3X 
-^  -    -J/^SV     .     /9R.X1    3R  /'/3R,'     ,     ,3R  = 
9F        V(-Jx)     +    (3Y-)'"9V-    V(g3f)     +    (§7) 
Using  these  in  equation  (1)  results  in  the  following  equation  for  range  error: 
where  the  notation  AP  =  AXP  and  AV  =  AX.  has  been  substituted. 
119 
(Reverse  Blank) 
(5) 
AR  -   |R  AP  +    |R   AV  (6)
DST-1000S-293-76 
12  March  1976 
APPENDIX  III 
CUTOFF  LAW  MECHANIZATIONS 
Appendix  II  established  the  concept  for  mechanization  of  thrust  termination.  The  desired 
condition  for  cutoff  is  when  the  ran^e  equation  goes  to  zero.  Setting  equation  (6),  Appendix  II,  to  zero 
results  in 
AR  «  0  =    -|y  AP  +   -|y  AV  (1) 
Then  equation  (1)  can  be  solved  for  the  required  cutoff  velocity  as: 
Vr  -    V„   -    Kp  (P   -    Po) 
where 
*"      aR/aV  (2) 
Equation  (2)  forms  the  basis  for  range  control  by  terminating  thrust  when  the  velocity  along  the 
principal  axis  equals  the  required  velocity. 
This  concept  for  range  control  is  simple,  requiring  only  two  scalar  quantities  (position  and 
velocity  along  their  respective  principal  axes)  and  equation  (2)  to  be  instrumented  in  order  to  satisfy 
the  range  control  function.  This  concept  relies  on  the  control  of  the  acceleration  vector  during  powered 
flight  to  be  of  sufficient  quality  to  satisfy  conditions  of  linearity  which  was  assumed  in  the 
development  of  equation  (2). 
To  implement  the  concept,  however,  requires  further  approximations  or  more  complex 
instrumentation  than  implied  above.  For  example,  the  acceleration  along  a  principal  axis  is 
a*  "  a.  cos  (6   —    Op)  —   gp  (3) 
where 
aP     •=  acceleration  along  the  principal  axis 
a.     «=  total  magnitude  of  sensed  acceleration  (sum  of  all  nonconservative  forces  +  mass) 
g  —  0P  a,  angle  between  a.  and  principal  axis 
gp     —  component  of  acceleration  due  to  gravity  along  the  principal  axis 
Integrating  equation  (3)  results  in 
VP  -  V.p  -  V* 
Pp  ~  P*>  —  P«p 
121
DST-10O0S-293-76 
12  March  1976 
w, here 
VP    «*  velocity  along  principal  axia 
Vip   ■»   integral  of  sensed  acceleration  along  principal  axis 
Vn.  «=  integral  of  gravity  along  principal  axis,  with  analogous  definitions  for  the  position 
(P)  terms. 
The  sensed  quantities  can  be  measured  by  using  inertial  instruments.  However,  the  gravity  terms  must 
be  computed  as  a  function  of  vehicle  position  which  is  not  easily  mechanized  in  principal  axes. 
If  it  is  assumed  that  the  flight  trajectory  is  sufficiently  close  to  the  reference  trajectory  that 
the  difference  in  the  gravity  integrals  from  nominal  is  small,  then  only  the  sensed  quantities  need  be 
measured.  The  error  which  results  from  the  assumption  is  termed  "mechanization  error."  The 
magnitude  of  this  contribution  to  the  total  mechanization  error  is  a  function  of  how  well  the  boost 
system  controls  the  flight  trajectory.  Therefore,  to  assess  the  magnitude  requires  a  detailed  analysis  of 
major  error  sources  contributing  to  trajectory  deviations  from  the  reference  trajectory.  Also,  given  one 
departure  from  the  theoretical  mechanization  (i.e.,  neglecting  explicit  calculations  of  gravity),  other 
than  theoretical  alignments  or  scaling  constants  (KP)  of  equation  (2)  could  result  in  smaller 
mechanization  errors. 
.  The  position  terms  in  equation  (2)  represent  differences  of  large  numbers  which  are  hard  to 
mechanize  accurately  without  resorting  to  digital  techniques.  An  alternate  approach  is  to  mechanize 
the  difference  between  the  reference  (programmed)  sensed  velocity  along  the  position  axis  and  that 
measured  by  an  integrating  accelerometer.  This  difference  is  small  (consistent  with  previous 
assumptions)  and,  therefore,  can  be  integrated  quite  accurately  without  resorting  to  digital 
integration.  With  this  consideration,  equation  (2)  is  modified  to  read 
where 
V™  -  V™  +  Kp  AP  (4) 
AP  -  j£*  
(VP  -  V.)  dt 
V™  «  required  sensed  velocity  along  velocity  axis  6  » 
Vo.  «  reference  sensed  velocity  along  that  axis 
VP  -  programmed  sensed  velocity  as  a  function  of  time  along  position  axis  6  P 
V.  -  measured  sensed  velocity  along  that  axis. 
122
APPENDIX  V 
(A)  SINGLE  ROTOR 
FTD  A75-1M 
(B)  DUAL  ROTOR 
Fig.  54    Soviet  Directional  Gyroscope* 
DST-1000S-293-76 
12  March   1976 
125
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12  March  1976 
APPENDIX  VI 
SOVIET  NAVIGATIONAL  AIDS  (U) 
1  ^^^-—--^(U)  Reference  to  an  inertial  navigation  system  will  occur  throughout  the  following  discussion.  It 
is  necessary  at  this  point  to  provide  a  statement  concerning  the  meaning  of  inertial  navigation  systems 
in  order  to  clarify  the  context  of  the  discussion.  All  inertial  navigation  systems  are  dead  reckoning 
systems.  A  dead  reckoning  system  depends  upon  continuous  operation  and  periodic  fixes  inserted  as 
updates  to  correct  its  inherent  errors  and  reset  the  system.  The  periodic  fixes  must  be  supplied  by 
systems  which  are  independent  of  measureable  drift,  platform  maneuvers,  and  other  disturbing 
,  factors.  No  one  fixing  system  meets  all  these  criteria.  Therefore,  any  inertial  system  must  have  several 
i  different  fixing  methods  in  order  to  function  accurately. 
127
D3T-1000S-293-76 
12  March   1976 
LORAN  positions  are  determined  from  the  difference  in  the  time  of  arrival  of  pulses 
transmitted  simultaneously  from  a  master  and  two  slave  stations.  Greatest  accuracy  is  achieved  when 
the  observer  is  on  a  line  which  bisects  the  angle  formed  by  the  two  base  lines  from  the  master  to  the 
slave  stations  and  decreases  as  the  distance  from  the  stations  increases. 
No  special  receiving  equipment  is  required,  but  automated  receivers  are  on  the  open  market 
which  would  remove  the  basic  operator  errors  involved  in  counting  dots  and  dashes. 
128
DST-1000S-2CJ-76 
12  March  1976 
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'"  nEEo  °  »'  
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USoK,  April   1966.  v  . 
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133 
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