Synopsis of Report
on Development of Conceptual Design
of an Artificial Earth Satellite
This document was signed by Sergey P. Korolev
on 25 September 1956. It is the detailed technical plan for the
'Object D,' the first Soviet satellite project. The program was
approved by a decree of the USSR Council of Ministers on 30 January
1956 and envisaged the launch of a heavy scientific satellite
in 1957 at the start of the International Geophysical Year. The
Object D program was a a direct result of Korolev and Mikhail
K. Tikhonravov's request to the government in May 1954 to launch
an artificial Earth satellite. <P>
Korolev's position at the time was:
- Chief Designer and Chief of the Experimental Design Bureau
No. 1 (OKB-1).
- The Decision of January 30, 1956, stipulates creation in 1957
- 1958 of a non-orientated artificial earth satellite on the basis
of a missile under development (Object D), having the following
- satellite weight 1,000 - 1,400 kg
- weight of scientific research hardware 200 - 300 kg
- first test launch of Object D scheduled for 1957.
- This report will discuss the basic results of development
of the conceptual design of a missile to be used as satellite
It should be noted that development of this Conceptual Design
had not been conducted by an accident: it is the result of all
prior work of the organizations that had taken part in development
of the RDD long-range missile. Operations of these organizations
included work on the turbopump rocket engines, control systems,
a satellite tracking complex, a ground equipment complex, and
gyroscopic instrumentation. A number of organizations of the USSR
Academy of Sciences also took part: the V. A. Steklov Applied
Mathematics Institute, the Institute of Automation and Telemechanics,
etc. First works of M. K. Tikhonravov and his team and their participation
in the draft plan of the artificial satellite are of a special
During recent 5 - 7 years operations with DD long-range missiles
have been conducted by the OKB and by departments of the Head
Scientific Research Institute with development of scientific and
research themes, and a number of RDD missiles of increasing range
have been built by effort of the whole industry. I am not going
to discuss these operations in detail, because everybody here
is familiar with these operations.
- 1. Basic Objectives of Explorations
- with the Help of the Satellite
- The program of comprehensive scientific explorations envisaged
to be carried out on board the first satellite is wide-ranging
1. Measurement of density, pressure, and ion composition of the
atmosphere at 200 to 500 km altitudes.
2. Investigations into the corpuscular radiation of the sun.
3. Measurement of the positive ion concentration along the orbit.
4. Measurement of the inherent electric charge.
5. Measurement of magnetic fields at 200 to 500 km altitudes.
6. Study of cosmic rays.
7. Study of UV and X-ray solar spectrum areas.
8. Studies of possibility of survival and life of animals during
long-term residence on board a spacecraft.
To accomplish all this, the satellite has to accommodate on-board
equipment of various types and for various functions for conducting
scientific research as follows:
- telemetry hardware for recording scientific data, having a
programmable device controlling conduct of measurements;
- a memory and a radio command line for sending commands from
the ground and for transmitting the data recorded during conduct
of the scientific research back to the ground for reception at
the ground stations when the satellite is orbiting over the territory
of the USSR.
- In addition to the above-mentioned objectives of scientific
research, launches of the first satellite will have to allow the
following first experimental data to be obtained. It will be necessary
in the future for development of an improved orientated satellite,
which will be designed for orbiting at much higher altitudes and
will have a much longer orbit life:
1) data on the character of movement of the satellite, its operation,
and accuracy of measurement of coordinates and tracking data;
2) data on the character of the satellite's movements with respect
to the center of gravity;
3) data on satellite braking in the atmosphere, bearing in mind
scarcity of our knowledge in this respect;
4) data on the thermal conditions of the satellite in orbit;
5) data on the power supply problems.
Those are in brief our objectives concerning the satellite.
The operations aimed at creating the first artificial earth satellite
represent, beyond any doubt, an important step in the way of mankind
into the universe, and we are now entering a new field of the
missile technology associated with development of the interplanetary
As a result of a thorough elaboration of the program of research
operations to be conducted on board the satellite, the Commission
of the Academy of Sciences chaired by Academic M. V. Keldysh has
found that one option of the satellite is not enough, and it has
been deemed reasonable to have three options with different sets
The weight of the satellite, based on components of equipment
and bearing in mind availability of the existent power supplies,
the radio telemetry system, tracking equipment, etc., is about
1,250 kg. This includes the weight of the shell of about 250 kg.
- 2. Specifics of the Satellite Design
- 1. Absolute tightness and air pressurizing to maintain a constant
2. Severe thermal conditions and the need of thermal control within
+5 to 30C (thus a temperature of 10 to 20C is required for operation
of the cosmic rays research hardware).
3. A large quantity of structural elements of equipment, modules,
mounting assemblies, etc.
4. Numerous pickups on board the satellite, each having its own
To insert a satellite of the necessary weight into orbit, it is
necessary and advantageous to modify operating conditions of the
propulsion unit of the central module by bringing them closer
to those optimal for a given product, based on the available power
data of the missile. It is assumed that the central propulsion
will be throttled down to about 60 tons of the pull beginning
with the lift-up moment. V. P. Glushko will give a more detailed
information on the experimental studies aimed at building the
- 3. Choice of the Orbit Parameters
- For these power conditions and for the missile parameters
for a given weight of the satellite, the satellite can be inserted
into different orbits. The choice of reasonable orbit parameters
was made first based on the need to achieve a long enough orbit
life (close the maximum), and second, based on the perigee altitudes
that are not too small (> 200 km). This is especially important
if the density of the atmosphere proves greater than expected.
The projects assumes the procedure of propulsion deactivation
by means of an integrator set up two times below the guaranteed
propellant reserve (with respect to the nominal reserve).
In this case, the propulsion will be deactivated by the integrator
for 90% of all launches, with the velocities at the end of the
active leg for the above-mentioned 90% of the products (7915 20
m/s) being 65 to 70 m/s higher than the velocity occurring with
the nominal guaranteed reserve. The rest of the products (10%)
will have a scatter of velocity within the above mentioned range
of 65 to 70 m/s (7850 to 7915 m/s).
This gives the following orbit parameters for two cases, respectively:
a) with deactivation by the integrator, using 50% guaranteed residues
Gguar. = 0.5nom.guar. (90% of
b) with deactivation after burning out propellant in the worst
case (corresponding to vn = 7850 m/s).
It should be noted that about 190 m/s are added owing to the earth
rotation during launch to northeast, taking into account the launch
point latitude (azimuth 35).
For each of these two cases, the nominal values of the orbit parameters
can be determined (in the event there is no scatter of the parameters
at the end of the active leg) ,and the limit values of the orbit
parameters can be determined (corresponding to the worst combination
of scatter of the parameters at the end of the active leg).
The ultimate parameters were calculated on the basis of the following
vn = 20 m/s; n = 0.6; hn = 6
It should be noted that the satellite life span values were calculated
based on the Mitre data on density of the atmosphere as recommended
by the GeoFIAN.
Based on some other data (e.g., according to Spitzer), density
of the atmosphere at 200 to 230 km altitudes is several times
as great in comparison with the Mitre data, and it is 10 to 100
times as great at the altitudes of 300 to 400 km. At the same
time, the object life span is approximately inversely proportional
to the density at the altitudes of 200 to 250 km.
For these reasons, the drag for the object in determining the
life-span was assumed to be two times as short as the calculated
time so as to have the upper limit value assessment, bearing in
mind a potential inaccuracy of the theoretical calculation of
aerodynamic coefficients at such altitudes. It will be required
to have a perigee altitude of at least 200 km.
A greater fraction of the reserve could be used, or the engines
could be even run without deactivation by the integrator, but
in such case the scatter of the orbit parameters would increase
(the scatter of the one revolution period is seven minutes for
the case of deactivation upon propellant burn-out).
- 4. Specifics of Separation of the Stages
- Throttling down the central propulsion impairs the separation
process and can result in a risk of collision of the separated
stages because of the relatively low acceleration values. This
problem is resolved by delaying separation until a high altitude
is reached and by throttling down a side-mounted propulsion (to
75% of the initial pull) about 17 seconds before separation. Throttling
down the side-mounted propulsion reduces the dynamic head at separation
from 145 to approximately 100 kg/m2, but it also results
in the velocity vn being decreased by about 15 m/s.
At the same time, throttling down the side-mounted propulsion
reduces loads during separation and allows the central object
propellant module to be retained.
Therefore, the main differences in the modified product are as
- the central propulsion pull is lowered to about 60 t (in the
vicinity of the earth); the side-mounted propulsion is throttled
down about 17 seconds before separation;
- the radio control hardware is removed (weight saving of about
- the radio module is replaced by an adapter module for attachment
of the product to the satellite;
- the rocket-based measurement system is minimized.
- With all the above modifications, the product can be launched
with a steady flight, the stages can be separated, and the satellite
with a preset weight can be inserted into an orbit with the errors
of ; n = 0.6; and vn = 20 m/s.
The pressurization value and the thickness of all load-bearing
shells remain the same.
- 5. Brief Characterization of the Orbit
- The satellite orbit will extend over a large area of the earth.
The flight altitude and the time of flight over the USSR, North
America and especially in the region of Mirny settlement for passage
over the region of the magnetic maximum are given in the Table.
- Satellite Altitudes and Flight Time over the Territory
of the USSR,
- People's Democratic Countries and North America
|Orbit revolution No.
||Total in 24 hours|
||16||163 minutes (11%)
|Flight over USSR and People'||Flight time, minutes
||17||163 minutes (11%)
||12 rev.||Total about 15 rev.
|--||Altitude, minimum, km
|--||Altitude, maximum, km
|Flight over North||Flight time, minutes
||124 minutes (8.5%)|
|--||Altitude, minimum, km
|--||Altitude, maximum, km
Total flight time over USSR over 24 hours is 137 minutes (9.5%).
Note: The data correspond to the nominal orbit during the first
24 hours and to the nominal propellant reserve.
6. Basic Problems in Satellite Design
Provision of the required temperature conditions on board the
satellite (0 to 30 and 10 to 20 for certain instruments).
On-board hardware power supply.
On-board hardware operation control (according to a preset timed
Provision of a radio telemetry system with a memory.
Provision of a tracking complex.
Sealing of the satellite for a prolonged period.
Provision of a system of omnidirectional antennas.
7. Specifics of the Thermal Conditions
The thermal conditions are characterized by material changes in
the thermal exposure factors: solar radiation, solar radiation
reflected by the earth, and substantial heat release from the
The components of the thermal balance sheet are as follows:
- direct solar radiation [about 1160 kcal/(m2hr)];
- solar radiation reflected by the earth (about 40% of the direct
- earth radiation;
- atmospheric air friction;
- heat of recombination of atomic oxygen on the satellite surface;
- heat release from operating on-board hardware (from 200 to
The thermal conditions are controlled by means of a radiation
wall of the sealed module, irradiating heat into space owing to
a high degree of opacity (>0.8) in the infrared spectrum area
( is the coefficient of opacity for the overall normal radiation).
A special coating of this wall assures low absorption of the solar
radiation (the coefficient of absorption As 0.3 for
the visible and ultraviolet areas of the spectrum, in which the
solar radiation energy has its peak value).
Transfer of internal heat release is assured by forced circulation
(by a fan) of nitrogen in the sealed module through a passage
adjacent to the radiation wall. When temperature decreases, this
passage is closed by a valve to cause a material reduction of
heat removal to the space environment. An additional thermal control
device is in the form of louvers on the radiation wall. Weight
of the thermal control system of the main sealed module is 60
to 70 kg together with power supplies.
Bearing in mind special requirements imposed upon cosmic ray research
hardware, a special thermostatic module is provided and isolated
from external exposure.
The sealed module surface will be protected on the insertion leg
against aerodynamic heating by means of a drop shield with panels.
The thermal conditions of the satellite on the launch pad will
be controlled by ground equipment because there are no weight
resources for an additional on-board device.
The above mentioned coating ( > 0.8; As 0.3) is
crucial for assuring the thermal conditions. It is necessary to
investigate its properties in orbit. Research in this area has
not been very extensive.
The calculation shows that preset thermal conditions can be realized
with the chosen layout of the satellite.
8. On-Board Hardware Power Supply
Power supply is assured by using electrochemical current sources:
silver-zinc storage batteries and mercury oxide batteries.
At the same time, the weight characteristics of the power supply
system are poor (a weight of up to 450 kg) and the operation time
is short. The reason is both a low power capacity of the batteries
(50 to 70 Whr per 1 kg on the average) and high energy consumption
of the on-board hardware.
It is necessary to expedite development of a solar array and to
work for lowering energy consumption of the hardware.
9. Experimental Debugging of the Satellite Design
1. Experimental debugging of functioning of all hardware and telemetry
2. Debugging of sealing, lead-outs, etc.
3. Experimental debugging of thermal control:
- building a full-scale thermal mock-up with real operating
- experimental investigations into heating of the satellite
structures in the insertion leg;
- experimental investigations into properties of special coatings
for the radiation surface.
The thermal mockup for studying the internal thermal conditions
will be tested in a special plant assuring the design temperature
of the sealed module shell, thus reproducing the internal thermal
conditions within the satellite.
The external thermal radiation exposure factors that determine
temperature of the shell can be calculated accurately enough.
4. The experiments aimed at studying properties of special coatings
in orbit are also crucial, bearing in mind high vacuum, collisions
with molecules and ions of rarefied gas at velocities greater
than 10 km/s, ultraviolet radiation of the sun, etc. These experiments
can be conducted by specialized institutions of the USSR Academy
5. Debugging the electrochemical power supply sources (hydrogen release and
Document obtained, edited, and translated by Dr. Asif Siddiqi
Dr. Steven Dick, NASA Cheif Historian
Steve Garber, NASA History Webmaster
For further information, please email firstname.lastname@example.org
Last Updated: August 1, 2007