An attitude control system is described in which angular rate signals are generated by rate gyros mounted closely adjacent to gimbaled engines at the rear of a vehicle. Error signals representative of a commanded change in vehicle angle or attitude are obtained from a precision inertial platform located in the nose region of the vehicle. The rate gyro derived signals dominate at high frequencies where dynamic effects become significant, and platform signals dominate at low frequencies where precision signals are required for a steady vehicle attitude. The blended signals are applied in a conventional manner to control the gimbaling of vehicle engines about control axes

Rupp, C. C.

Vonpragenau, G. L.

English

NASA Technical Reports Server

United States Patent [191 
von Pragenau et al. 
[ 1 1  I B 3,984,072 
1451 Oct. 5 ,  1976 
ATTITUDE CONTROL SYSTEM 
Inventors: 
Assignee: 
Filed: 
Appl. No.: 
Published 
George L. von Pragenau; Charles C. 
Rupp, both of Huntsville, Ala. 
The United States of America as 
represented by the Administrator of 
the National Aeronautics and Space 
Administration, Washington, D.C. 
Oct. 2, 1974 . 
5 1 1,346 
under the second Trial Voluntary 
Protest Program on January 27, 1976 as 
document No. B 5 1 1,346. 
US. CI. ............................... 244/169; 24413.2 1 ; 
2441 17 1 
Int. CL2 ............................................ B64G 1/20 
Field of Search ........... 24411 64, 165, 17 I ,  776, 
244155, 169, 3.2, 3.21; 2351150.2, 150.25 
References Cited 
UNITED STATES PATENTS 
3.1 32,828 511964 Edinger e t  a]. .................. 244/3.2 X 
3,301,508 1/1967 Yamron ........................... 24413.2 X 
3,395,874 8/1968 Bartoli et a1 .......................... 244/32 
COM PUT E R 
Y b  
FOREIGN PATENTS OR APPLICATIONS 
952,261 311964 United Kingdom .................. 24413.2 
Primary Exuminer-Trygve M. Blix 
Assistunt Examiner-Barry L. Kelmachter 
Attorney, Agent, or Firm-L. D. Wofford, Jr.; G .  J. 
Porter; John R. Manning 
[571 ABSTRACT 
An attitude control system for a space vehicle in 
which angular rate signals are generated by rate gyros 
mounted closely adjacent to gimbaled engines at the 
rear of a vehicle, and wherein error signals representa- 
tive of a commanded change in vehicle angle or atti- 
tude are obtained from a precision inertial platform 
located in the nose region of the vehicle. The rate 
gyro derived signals dominate at high frequencies 
where dynamic effects become significant, and plat- 
form signals dominate at low frequencies where preci- 
sion signals are required for a steady vehicle attitude. 
The blended signals are applied in a conventional 
manner to control the gimbaling of vehicle engines 
about control axes. 
1 Claim, 2 Drawing Figures 
Y'Y 
321 
S U M M E R  
2 
 
,"'Y 1'y + d y  
https://ntrs.nasa.gov/search.jsp?R=19770003171 2020-03-20T06:41:20+00:00Z
U.S. Patent oct. 5,  1976 3,984,072 
3,984,072 
1 
A"T1TUDE CONTROL SYSTEM 
ORIGIN OF THE INVENTION 
The invention described herein was made by employ- 
ees of the United States Government and may be man- 
ufactured and used by or for the Government of the 
United States of America for governmental purposes 
without the payment of any royalties thereon or there- 
for. 
BACKGROUND OF THE INVENTION 
I .  Field of the Invention 
This invention relates to directional control systems, 
and particularly to an attitude control system for a 
space vehicle. 
2. General Description of the Prior Art 
As space vehicles have been made larger, the prob- 
lems associated with controlling their flight characteris- 
tics have vastly increased. In order to provide the 
needed sophistication in control systems, delicate and 
bulky apparatuses are required, and this often necessi- 
tates locating the control sensors spaced substantially 
from the engines, typically locating the control sensors 
in the nose region of the vehicle. Since a space vehicle 
is not a rigid object, vibrations and resonances develop, 
wherein movement of one portion of the vehicle is 
typically not in phase with movement of another por- 
tion of the vehicle. As a result, the attitude of the vehi- 
cle at one location is often not the same as the attitude 
elsewhere on the vehicle. The magnitude of this differ- 
ence becomes particularly large when the attitude of 
the vehicle at the nose of the vehicle is instantaneously 
compared with the attitude of the vehicle at the tail of 
the vehicle. The difference reflects bending moments 
which exist along the length of the vehicle and complex 
compensations must typically be applied to control 
signals derived from attitude measurements in the nose 
of the vehicle to produce optimum commands to gim- 
baled engines located at the tail of the vehicle. The 
compensations are determined from extensive ground 
vibration tests of the vehicle and from elaborate com- 
putations which take into account propellant deple- 
tions and payloads for h particular flight. Not infre- 
quently, however, after initial tests and computations 
are made, new flight plans are developed, necessitating 
new tests and computations. The result is a tremen- 
dously expensive effort in determining answers to stabi- 
lization questions. 
OBJECT OF THE INVENTION 
5 
10 
15 
20 
25 
30 
35 
40 
45 
50 
It is the object of the present invention to provide 
what might be termed a less critical attitude control 
system which will not be as sensitive to vehicle vibra- 
tions and resonances. enabling less compensation of 55 
attitude control systems for various vehicie configura- 
tions. 
SUMMARY OF THE INVENTION 
In accordance with the invention, strapped-down or 
fixed-mounted angular velocity sensors are positioned 
in the tail region of the vehicle near gimbaled engines 
and near where the vehicle is actually controlled. Typi- 
cally, such devices would be rate gyros, and there 
would be one for the measurement of angular rate 
about each of the body coordinate axes of the vehicle, 
the roll, pitch, and yaw axes, A precision sensor such as 
a gyro stabilized platform would be conventionally 
positioned remote from vehicle engines, typically in the 
nose of the vehicle. Associated with it would be a guid- 
60 
65 
2 
ance computer which would determine attitude com- 
mands for each of the three axes to cause the vehicle to 
assume commanded attitudes in order to perform se- 
lected maneuvers at particular times. Corresponding 
angular axis rate signals from rate gyros would be 
blended with each axis guidance signal as follows. 
Each discrete axis command signal from the guid- 
ance computer would be filtered by a low-pass filter to 
block signals of a frequency higher than the lowest 
resonant frequency of the vehicle. Each rate gyro oper- 
ates as a differentiator of the rotation angle of the rate 
gyro with respect to inertial sp,ace and provides a signal 
output having an amplitude characteristic which rises 
in amplitude with frequency and provides little output 
below the lowest vehicle resonance f,. The resulting 
differentiation signal is passed through a low-pass filter, 
as in the case of the error signal and the inertial plat- 
form, and this filter produces an integrating effect on 
the, frequencies passed, being, of course, a band of 
frequencies just abovefo. In this way, signals are derived 
from the rate gyro which are in an adjoining and thus 
complementary range of those just below fo from the 
inertial platform. And, significantly, the rate gyro derived 
signals are only present during relatively rapid vehicle 
movement. 
The guidance computer signals and rate gyro derived 
signals are summed to provide a frequency dependent 
blend of the two signals. Alternately, guidance com- 
puter signals and rate gyro signals would be summed 
and then subjected to a low-pass filter. Thus, with ei- 
ther arrangement, when the space vehicle is relatively 
steady and signals are below the lowest vehicle reso- 
nance, the output of the guidance computer will domi- 
nate the control signal; and when the vehicle move- 
ment contains frequencies above the lowest vehicle 
resonant frequency, the angular error signals derived 
from the rate gyro will dominate. In such instance, the 
latter signal would be superimposed on a guidance 
computer derived signal to provide an instantaneously 
accurate and up-dated error control signal. derived 
from the rate gyro which are in an adjoining and thus 
accurate and up-dated attitude error control signal for the 
rate gyro location. 
The thus obtained control signal is typically applied, 
or further compensated and then applied, to gimbal 
control devices which appropriately adjust the angular 
orientations of rocket engines to effect a desired stabili- 
zation or change of vehicle attitude. 
BRIEF DESCRIPTION OF THE DRAWINGS 
FIG. 1 is a schematic illustration of an embodiment 
FIG. 2 is a schematic illustration of a portion of an 
of the invention. 
alternate embodiment of the invention. 
DESCRIPTION OF THE PREFERRRED 
EMBODIMENT 
Referring to the drawings, space vehcile 10 is gener- 
ally depicted in simplified form. Stable or inertial plat- 
form 12 and guidance computer 14 would be conven- 
tionally interconnected and packaged together near the 
nose of the vehicle, as shown. The stable platform pro- 
vides a fixed inertial reference and the guidance com- 
puter delivers and provides three outputs, each repre- 
sentative of one of the angles by which it is desired to 
rotate the vehicle to perform a desired maneuver at 
attitude correction. These are identified as a,, a,, and 
a, and related, respectively, tot the roll, pitch, and yaw 
angles. They are computed in terms of vehicle orienta- 
3,984,072 
3 4 
tion as seen at the nose of the vehicle with respect to to a vehicle and which provide an immediate output in 
axes 16 of stable platform 12. terms of body coordinates. 
Rate gyros 18 are each mounted on fixed mounting An alternate arrangement of the invention is illus- 
plate 19. Instead of being located in the same location trated in FIG. 2 and in it the function of low-pass filters 
with the inertial platform and guidance computer (as 5 28 and 30 are combined in a single filter 36, it being 
has been the case), mounting is in the tail section ofthe located after summer 32. In such case, there would be 
vehicle relatively near gimbaled engines 20 and 21. provided appropriate level controls 38 and 40, one for 
Three fixed-mounted or "strapped-down" rate gyros each channel preceding an input to summer 32. 
18 are attached to mounting plate 19, and thus the rate If there is only one gimbaled engine, roll control may 
movement of the vehicle about one of vehicle axes 22, tion of a roll rate gyro. Pitch and Yaw would still be 
as they appear at rear-mounted gyro mounting plate 19 controlled by gimbaling of the one engine. 
fid as d&/dt, db /d t ,  and db,/dt for the roll, pitch, and of signal delays encountered with precision guidance 
yaw axes, respectively. 1 5 computers, auxiliary means have been employed to 
gyros provide outputs representative of rate of angular 10 utilize reaction control jets in the vicinity of the loca- 
with respect to inertial space. These outputs are identi- It is to be noted that in order to alleviate the problem 
Since vehicle processing is generally the Same for 
each signal and control function, only one channel of 
control (the pitch axis channel) is illustrated in detail. 
Provide approximate command control signals in the 
interim between normal outputs. BY use Of the present 
invention, supplemental signals need not be generated- 
The angular pitch rotation b, of mounting plate 19 It is acknowledged that in the past some effort has 
(with respect to inertial space) is applied to rate gyro 20 been made to derive attitude signal information from 
is passed through low-pass filter 28. The combined inaccuracies due to biasing and other cumulative gyro 
effects are to provide a band of position information errors. In the present case, these errors are avoided 
since gyro signal use is limited to short periods of time signals above the lowest resonant frequency of vehicle 
10, labeled output br,. This output represents an incre- It is to be further appreciated that in certain dynamic ment of attitude error that is a pitch angle signal which situations several strapped-down sensors may be used, at any instance is equal to  the increase or decrease in particularly to provide outputs at particular frequen- 
cies of vibration and at particular places on the vehicle. attitude error that has occurred. 
The pitch command Output from guidance 'Om- 30 Such applications are quite feasible because of the low 
described would then be applied for at least one 
to Summer 32 where they are 35 down sensors. It is also feasible to arrange an updating 
18, and the responsive electrical output of rate gyro 18 rate gyros by integration. This, however* has lead to 
25 when critically needed. 
puter l4 is to low-pass 309 which cost of strapped-down sensors. The blending technique 
frequency to provide a thus 
Output "U Of 
frequencies higher than the lowest vehicle resonant 
Output "U' The 
30 and the Output b'U Of low-pass 
strapped-down sensor, and only one precision Sensor 
would be required to update one or several strapped- 
added. The sum a'u+b'U is a chain where a first strapped-down sensor would be 
control 34, which in turn Operates 
A like system is employed for control Of the Other 
28 are 
pitch angular error signal, and it is fed to pitch gimbal updated by a precision sensor, and the combined sig- 
and 21 to effect attitude control of the vehicle in pitch. second strapped-down and so forth. such 
of control of the vehicle. time constant of each would be progressively lower. 
The signal blending resulting from the Processing of In summary, the present invention permits a free 
the signals as described above enables a control signal choice of precision location independent of 
to be extremely accurately generated for both the sta- vehicle dynamics. The blending technique updates 
tionary case Of the vehicle where Oscillatory frequen- 45 strapped-down rate Sensor signals and reduces the 
ties f are less than Or equal to f o ,  the frequency of the computational effort because the signals from the 
lowest vehicle resonant, and for the dynamic Case strapped-down device are given in the required body 
where vehicle ~ ~ ~ n a n c e S  f are greater than fo. Thus, fixed coordinates, and the updating by the precision 
the present invention enables dominance Of a Control signal can be performed at a slower rate. Vehicle con- 
system by the % signal from guidance computer 14 50 trol is corrected by low cost strapped-down sensors, 
when the frequency f 1s less than or equal to f o  and sensing vehicle attitude information at essentially the 
enables the system to be dominated by the signal de- point where control forces are applied. This enables a 
rived from the rotation angle b, ofthe rate gyro mount- signal accuracy over a wide range of vehicle vibrations 
ing plate, typically for short periods, when the fie- which approaches the obtainable from a precision iner- 
quency f is greater than fo. For short periods, the 55 tial system operating in a more limited range. Bias and 
mounting plate angle (with respect to inertial space) other errors common to strapped-down sensors are not 
near engines 20 and 21 produces an extremely accurate integrated on a long-term basis, and thus accuracy does 
control signal that can be applied to engines 20 and 21 not deteriorate with time. 
for its control. Thus, at a time when oscillations are What is claimed is: 
above the frequency fo and are particularly critical, 60 1. In an attitude control system for a space vehicle 
they are rapidly taken into account. Usual computa- including at least one aft mounted and gimbaled motor, 
tional delays inherent in obtaining signals from a typi- a motor gimbal control, the guidance means positioned 
cal guidance computer are avoided. Such a computer in the forward region of said vehicle having an inertial 
normally digitally calculates angular error signals in reference platform and computing means for comput- 
terms of an inertial or fixed coordinate system and then 65 ing and providing attitude error signals in terms of a 
translates the signals to vehicle body coordinates be- fixed inertial set of axes, the improvement comprising: 
fore the signals can be used, and this takes a significant angular rate sensing means comprising at least one 
time. This is not true of rate gyros which are strapped rate gyro fixedly mounted on said vehicle in the aft 
motors 2o nals from a first blender would be used to update a 
40 case, successive filters would be employed wherein the 
3,984.072 
5 
end region of said vehicle for providing at least one 
set of rate signals representative of the rate of rota- 
tion of said vehicle about at least one axis, said rate 
signals increasing in amplitude with increasing fre- 
quency and having a relatively low output at fre- 5 
quencies below the lowest resonance of said vehi- 
cle; 
summing means responsive to attitude error signals 
from said guidance means and rate signals from 
said angular rate sensing means for summing said 10 
rate signals and said error signals; 
low-pass filter means responsive to summed signals 
from said summing means for applying an integrat- 
ing and attenuating effect to said summed signals 
15 
6 
above the lowest resonant frequency of said vehicle 
and providing control signals to said motor 
gimbal control; and 
whereby for conditions of absence of vibrational 
frequencies higher than said lowest vehicle reso- 
nance, said control signals are dominated by atti- 
tude error filtered signals, from said guidance 
means, and during the occurrence of vibrational 
frequencies of a freqency higher than said lowest 
vehicle resonant frequency, said control signals are 
dominated by filtered signals derived from said rate 
sensing means. * * * * *  
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65 


Attitude control system

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