JPRS ID: 9303 USSR REPORT SPACE

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~ C
~E~TE1~~E~ ~~~0 ~ F~U~ ~ i

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- JPRS L/9303
17 September 1980
USSR Re ort
~
SPACE
(FOUO 8~/80)
,
FBIS FOREIGN BROADCAST INFORMATION SERVIGE
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NO i'E _
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JPRS L/9303
17 5eptemher 1980
USSR REPORT
SPACE
(FOUO 8/80)
CONTENTS
_ LIFE SCIENCES
Soyuz-T Survival Suits Pictured 1
- SPACE APPLICATIONS
Trajec~ories of Space Vehicles Intended for
Investigation of the Earth From Space 2
' ~ Meteorological Sounding of the Sub3acent
Surface From Space 12 -
Materials Processing in Space....... 14
Aerospace Methods of Studying Soils 20
The Study of Taiga Topography by Remote Methoes 26 ~
SPACE EPIGINEERING
System for the Emergency Rescue of Cosmonauts From
Orbital Stations [SASKOS] 28
On Aperture Synthesis Using a Space Rad~otelescope........... 36
On Selecting Types of Launch Vehicles for the
Implementation of a Space Research Program Over _
a P?inimum Period 'of Time 52
SPACE POLICY AND ADMINISTRATION
Soviet Perceptions of United States Remote
Sensing Program 65
- a- jIII - USSR - 21L S&T FOUOJ
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LIFE SCIENCES
SOYUZ-T SURVIVAL SUITS PICTURED
Yaris AIR & COSMOS in French No 819, 28 Jun 80 pp 44, 56
[Article by S.B.: "'Soyuz-T', A New Generation Soviet Vehicle"]
[Excerpt] A new space suit, called a"survival suit," was used [on board -
Soyuz-T]. It protects cosmonauts in the case of acute cabin depressuriza-
tion. This suit is lighter (two times lighter than that worn by Gagarin),
more flexible, and can be put on rapidly. Thanks to a"transparent steel"
visor, the cosmonaut's field of vision is greater. Even when the suit is
inflaCed, the cosmonaut can, while *aearing his gloves, wind his watch cr
pick up a needle.
- "wr* ~ ~'~,'~,~-,y ~t,~ ~ ~ z_ p,
; ,~.�,z,~, x~.~y? ~ yc~/-, a4~� - .
d .~h. .M 4 y~
~ ~ t ,~3 f~ 1~."
. y,'..; ,~.ti Yf
` 4~
'ri .
~ K I , , k ~
~ 1
Y'n., ~~4 ~'r, . ~1.
. . ~
~ I~ . . F V ' ~
. ~t- . ~ Y ka~, t ~,"j� ! ~lf/
~ ~
. .r~- ~ ~ r
, ~C~. , , ~ ~ ~ ` r , -
~
' ~
� ~~LJ ~
t ~ ~
The Soyuz-T crew, i~falyshev and Aksenov (left to right), wearing the new
space "sur.vival suit" which was flight tested for the first time.
COPYRIGHT: A. & C., 1980
[1853/11-P]
CSO: 1853
- - 1 -
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SPACE APPLICATIONS
UDC 629.19:551
TRAJECTORIES OF SPACE VEHICLES INTENDED FOR INVESTIGATION OF THE EARTH
FROM SPACE
Moscow ISSLEDOVANIYE ZEMLI IZ KOSMOSA in Russian No 3, 1980 pp 91-97
[Article by Ye. L. Lukashevich, State Scien~tific Research and Production
Center "Priroda"]
[Text] One of the important stages in investigating the earth from space
is the collection aboard a space vehicle of information concerning the
,~J state of the atmosphere, lithosphere and hydrosphere of our planet. The
_ realization of this stage involves, in particular, solving the problem of
positioning the space vehicle in near-earth space in such a way as will
allow an arderly scanning of the entire surface of the planet or a stipu-
lated spherical zone in a relatively short time interval, taking into
account a number of specific requirements and restrictions on the differ-
ence in scale of the collected information, on the periodicity ~f repeated
observations of one and the same regions, on the illamination of the earth's
surface, etc. We will examine the peculiarities of the trajectory of a
space v~hicle which will satisfy the most important of the enumerated
requirements as well as methods for computing the nominal parameters of
these trajectories.
A necessary and adequate condition for ensuring a full scanning of the
earth's surface within the limits of the observed latitude zone is that
the orbit belong to a class of quasigeosynchronous trajectories with such
a relationship between the Draconian period of revolution of the space
vehicle and the diurnal rotation of the earth in which the trajectory of
the K-th revolution will pass in the neighborhood of the trajectory of
- the (K - Nm)-th revolution. In this case the number of revolutions Nm
between two successive passes of the trajectory through the neighborhood
of an arbitrary point on the surface is the "order" of the orbit and the
number of days m in which the space vehicle makes Nm revolutions is the
"multiplicity" of the orbit. The distance between the trajectories of the
(K - Nm)-th and K-th revolutions is 3etermined by the width of the zone
of scanning of the on-board observation systems.
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There are different algorithms for computing the nominal parameters of
the quasigeosynchronous trajectories, differing from one a!nother in the
use of different intermediate orbits. However, all these algorithms are
derived from an expression representing the mathematical writing of the
principle of quasigeosynchronicity of mation:
~L-2nm- TNm-f-~SZm, (1)
where L~ L is the angular displacement of the flight trajectory in Nm r2vo-
lutions, t.J is the angular velocity of the earth's diur.nal rotation, T is
the Draconian period of revolution of the space vehicle, 0,~ is the change
in longitude of the ascending node in m days. We can become familiar with
a general case of computation of geosynchronous trajectories, constructed
in Keplerjan intermediate or'bits, in such studies as [1]. Applicable to
the problem of investigation of natural resources the algorithm for comput-
ing the altitude of a circular quasigeosynchronous perturbed Keplerian or-
bit was proposed in [2].
We note that in most cases the Q~ m and L~ L parameters are commensurable
and therefore the choice of parameters and the evaluation of stability of
quasigeosynchronous trajectories must be carried out with rigorous allow-
ance fo~ precession of the orbital plane, caused primarily by the influence
of acentrality of the planetary gravitational field. In this connection it
is desirable in the stage of ballistic planning of such trajectories to use
models of the gravitational field taking into account the fundamental har-
monics of the earth's gravitational potential. As such a model in [3] use
was made of the gravitational field of three pairs of fixed centers with
complexly conjugate masses situa.ted on the z~-axis of a geocentric equator-
ial coordinate system. According to the estimates cited in [4], such a
model of the gravitational field makes it possible to take into account
the perturbations caused by the influence of the totality of the f irst
2ona1 harmonics the earth's potential (to the eighth inclusive).
Now we will examine the procedure for computing the nominal parameters of
quasigeosynchronous orbits based on the simplified algorithms in [3]. Ex-
panding the T and Q~ parameters in expressicn (1), after simple trans-
. formations it is easy to obtain an equation transcendental relative to
the semimajor axis which with an accuracy ta terms of the order of the
square of the earth's flattening inclusive can be represented in the fol-
lowing form: _ . _ .
v~M eL, s 3 3 ll
a~ ~ _ ~Nm ~ ~ _ 2 aEOZNm ~ r q Z e�Z (1 - y sZ+eo aZ / J '
L ~2~
where f is the gravitational constant, s and oc are the sine and cosine of
the angle of inclination, Ep~ R~/ap(1 - e02), M and R are the mass
and mean equatorial radius of the earth, I2 is a coefficient on the second
zonal harmonic of the earth's gra~ita~ional potential, This expression
makes,it possible to use the iteration method to determine the nominal
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rux urrl~l~. ua~ vtvLx
semimajor axis ap of a ciuasigeosynchrcnous orbit using the stipulated val-
ues of the initial eccentricity ep, inclination i, initial displacement -
of the trajectory 0 Lp and the range of the working altitudes [hmin, hmax~~
which characterizes the order Nm a.nd multiplicity m of the orbit.
'rvEs note that from the condition of obtaining information with minimum
scale scatters it is preferable to have circular orbits for the considered
class of space vehicles. However, as a result of the peculiarities of putt-
ing a space vehicle into an orbit with the mean altitude h~[hmin~ hmax~
as a rule one obtains a near-circular trajectory with a certain e0 value.
The inclination of the orbital plane to the equatorial plane is selected
from th~ condition i=~f'~X, where ~~X is the maximum latitude ~of the
subsatellite point. Finally, the displacement of the trajectory DLp is
chosen so that it is commensurable with the width ~,of the scanning zone,
taking into account the necessary transverse overlap of adjacent zones at
the stipulated latitude It must be remembered that sign L~Lp deter-
mines the direction of the displacement: with ~ Lp ~0 the tra~ectory is
displaced to the east; with Q Lp ~ 0-- to the west.
The following sequence is proposed for computing the semimajor axis with
the use of equation (2).
1. Taking into account the possibility of putting a space vEhi.:le into
a working orbit, the width of the scanning 'zone when making observations
from a stipulated range of working altitudes, the lateral ove*_-lap of the
scanning zones at a stipulated latitude and the maximum latitude of the
territory sub~ect to observation, we will determine the elements ep,~d,LpJ,
i. In this case it is possible to use the obvious expression
- .
~OI.~~=(l'-,~1,')/Rcosc~~ - ~3~
where L* _,Q(s2 - oC2tg2t~)-1~2 is the extent of the scanning zone along
the parallel at the latitude ~!p* is the initial lateral overlap of
the scanning zones at the latitude cP .
2. Using the simplified expression following from condition (1)
c~ hm~o+ha,.= ~r~
4L~2n[m--yiM ~ 2 ) Nm], ~4~
we will determine the multiplicity and order of the orbit, as well as
sign !1 L. It must be remembered that m and I1~ are whole numbers related
by the expressions: N~ = 16m - n, where n= 1, 3,..., when m and n are
not multiple numbers and Nm~n - 1 for a case of multiple m and n. The
sought-for parameters must ensure min~F(Nm, r,~, sign Q L)~ , where F is
the difference of the right and left sides of the approximate equation
(4). Good results are obtained from a gr~~phic solution of equation (4) with
use of the preconstructed regularities ~ L= L1 L(h), where h=(h~in + hmax)
/2, for same m and Nm values. For example, in the range of worki_ng altitudes
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2d0-1400 km it is sufficient for this to have orbits having the foilow-
ing order characteristics: N1 = 16, N3 = 47, N2 = 31, N3 = 46, N1 = 15,
N= 44 N2 = 29, N3 = 43, N1 = 14, N3 = 41, N2 = 27, N3 = 40, N1 a 13,
3 ~
N3 = 38, N2 = 25.
3. Assuming that in a zero approximation a~~) = R+(h~x + h~in)/2 and
substituting the determine.~ ep, m, I~, dL~~values and the known para-
meters ~ p, s, cL' into equation (2), we will compute the semimajor axis
a01) in the first approximation.
4. Refining the � 0 parameter with al~owance for a~l), from expression (2)
- we obtain a second approximation of the semima~or axis a~2), Then, contin-
uing the it eration process, it is possible to determine the ap value with
any stipula ted accuracy.
~ It can happen that orbital eccentricity cannot be ascertained prior to the
determj.nation of ap. Then it is necessary that the successive approxitna-
tion in computation of the semima~or axis begin with the assumption that
in the zero approximation e~~) = 0. The radius rcir o~ the circular orbit
obtained under this conditi~n from equation (2) is assumed to be the semi-
major axis and the eccentricity is then determined from the minimum and
maximum altitudes which the carrier-rocket can ensure when puttin a
space vehicl e into an orbit with a0 = r~r. The eccentricities eal~ deter-
mined in such a way are agai.n introduced into expression (2) and we will
repeat the' entire procedure of computations for the purpose of refinin~ ap.
It is easy to note that by extending the region of quasigeosynchronous
trajectories to multiple orbits with m> 1, with virtually no limit we
will broaden the range of altitudes applicable for selecting the working
orbits of space vehicles of the considered class. The sequence for scann-
ing zones on the earth's surface from orbits of different multiplicity is
different. It can be shown that there is the following regularity making
it possible to estimate the total number of revolutions N~ necessary for
total scanning of the observed surface:
- - - . - - _
N~=[~lm]Nm-f-(~-m~'~~m~) ~I^(/m]+1)-9,
where y= TO Lp is the number of scanning bands in each interval between
revolutions, Tp is the initial value of the period of revolution. By insig-
nificantly c hanging the L~LD value it can be achieved that Y is a whole
number.
One of the interesting peculiarities o� use of multiple orbits is the pos-
sibility of selecting different observation schemes when working with one
and the same altitude or with near-lying altitudes. For example, a space
vehicle moving at an altitude N 390 km and having a scanning zone cf about
5� can have two order characteristics of a circular polar orbit: N2 = 31 ~
with displac ement of the flight trajectory in an easterly direction and
N3 = 47 with displacement of the tra~ectory in a westerly direction.
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~~~~K ~~~~i~ i~.ini, Iltih; l?Nl,v
Now we ~~ill examine the problem of illumination of features observed at
the planetary surface. This problem is one of the most important in the
planning and organization of space surveys with the use of photographic and -
television remote sounding systems.
The illumination of features situated on the earth's surface and in the -
neighborhood of the flight trajectory is determined by the relative ori-
entation of the orbital plane and the plane passing through the mean ter-
minator line. The principal factor governing precession of the orbital
plane is acentrality of the gravitational field. The secular component
of this precession in m days with an accuracy to the sq~uire of the earth's
flattening can be represented in the form
~]:m=-3nae Nm� `J~
The plane of the terminator also rotates as a result of the annual revolu-
tion of the earth around the sun. If oC> 0(i~ 90�), both planes rotate '
in opposite directions and the branch of the flight trajectory on which
the survey occurs periodically emerges from the region of solar altitudes
admissible for space photography. If the period of active operation of the
space vehicle does not exceed a month, the choice of the corresponding mo-
ment uf the launching can be achieved by continuous observation of the
earth with satisfactory illumination of its surface. However, if the per-
iod of active operation of the space vehicle is more than a month, with
emergence of the working branch of the trajectory from the region with
satisfactory illumination a pause appears in the operation of the space
vehicle and the continuity of observation is disrupted. With u< 0(i >
90�) both the orbital plane and the plane of the terminator rotate in one
and the same direction. Thus, it is posaible to select such an inclination
that the angular velocities of rotation of both planes will be equal and
heliosynchronous and the orbit will be solar--sqnchronous. The principal
advantage of space vehicle motion in such an orbit is a constancy of the
local time of transit over regions situated at one and the same latitude.
In addition, solar-sS~nchronous osbits have a number of inerits favoring an
increase in the effectiveness of long-term observation of the planetary
surface with the use of space technology.
The methods for computing the elements of a solar-synchronous orbit, espec-
ially with the use of Keplerian intermediate orbits, are we11 known. Here
we will use a somewhat different approaeh to choice of orbital parameters,
and in particular, inclination, based on the fact that a solar-synchronous
trajectory must at the same time be quasigeosynchronous.
Under the condition that in the course of one tropical. year T y the orbital
precession is a value 2 n, from expression (5), with an accuracy to terms
of the order of the square of planetary flattening inclusive, we obtain
the following equation transcendental relative to the nominal ap value
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av, - 3asa= ti ~i11 r1+ 3 e: (1- ~ sZ-~e,=a=
- a - - --2n' L 2 ~ l
(6)
where n* = 2?1/T r is the mean angular velocity of the earth's motion
around the sun. By comparing expressions (2) and (6) it is easy to derive
a simple expression for de~ermining the a value of a solar-synchronous
quasigeosqnchYonous orbit in dependence on d LQ, Nm and m:
2 2nm-OL,
a = - 3e~'Nm c~Ti-2n ~ (7)
For more rigorous computation of the orbital parameters it is necessary
to have a joint solution of equations (2) and (6) by the iterations meth-
od. As a first approximation of o~ it is possible to use a value computed
using formula (7).
Now we will turn to the problem of the stability of a quasigeosynchronous
orbit and the maintenance of its parameters, in particular, the semimajor
axis, in an admissible range. The principal perturbing factor exerting an
influence on the secular change in a0 and ep is the earth's atmosphere. An
evaluation of these changes in the ballistic planning stage for orbits with
a small eccentricity can be carried out using a static atmosphere using the
formula3 in [5J:
Da=-4ncao pn eap(-v) (1-I-v'/4),
De=-2nca, (1-e,=) p,~ eap (-v)v (1-~-v'/8),
~
where 0 a and L~e are the secular p~rturbation of the semima~or axis and
eccentricity during one revolution of the space vehicle; c is the ballis-
tic coefficient; /~rt is atmospheric density at orbital perigee; y/ = apep/H,
H is the height of the homogeneous atmosphere. In d days the semima,jor axis
changes by the value L~ adN~, which leads to a change in the Draconian
period of revolution and in the last analysis to a noncorrespondence be-
tween the displacement of the tra~ectory ~ L and the width Q of the scann-
ing zone. The period of revolution on the d-th day of flight with an ac-
curacy to values of the order E 2 inclusive is determined using the for- -
mula
T_ 2n ya' r 1_ 3 e: ( q_ 3 sa~-ezat I J'
y
fM ~ z ~ z .
where a= a0 + d adNm, e= e~ + d edNm. Disruptions in the conditions for
lateral overlaps of adjacent scanning zones are evaluated using expressions
(1) and (3). The latter makes it possible, on the basis of the L,L value,
changing after d days of flight, to determine the new d,Q* value at the
stipulated latitude For orbits with a w~sterly displacement of the
flight tra,jectories this disruption will be expressed~in the form of a pro-
gressive slipping of the zones relative to one another and this can lead to
an incomplete scanning of the inter-revolution intervals if the number of
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days D of active operation of the space vehicle is limited: D< L/~ L0,
where L is the angular value of the inter-revolution interval. For orbits _
with an easterly displacemsnt of the flight trajectories the disruption of -
ordered observation is expressed in a progressive spreading apart of the
scanning zones, which at a stipulated latitude SP does not make it possible
to obtain an integral picture when investigating features whose dimensions
exceed the value. For solar-synchronous orbits a decrease in ap leads,
in addition, to a disruption of the heliosynchronicity of precession of
the orbital plane.
Thus, the admissible deviations of the semimajor axi~ from the nominal val-
ue are determined by the range of the accepta3le values, for example,
0.03~ ~,~*~.~