The system of means of launching spacecraft to
geostationary orbit
Composition and parameters
About the prospects of the proposed launch vehicle
Competitiveness and value of the system
Composition: |
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A launch vehicle that includes a universal |
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upper stage. |
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Cargo flow on GSO, t/year |
22.5 |
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Productivity, g/s |
3.6 |
t / year |
67.5 |
Area of solar panels, m² |
670 |
Radiator area, m² |
57 |
Power of the electrolysis plant, kW |
92.5 |
Cryogenic plant power, kW |
7.5 |
Carrier aircraft: special modification of the AN-124 or more |
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light aircraft. |
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Rocket block: |
0.035 |
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Starting weight, t |
69.0 |
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Payload mass, t |
3.1 |
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Number of launches per year |
30 |
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Fuel mass for the main propulsion system, t: |
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liquid oxygen |
51.2 |
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liquid hydrogen |
5.8 |
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êåðîñèíà |
5.2 |
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Specific impulse of the propulsion system, N*s/kg: |
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in three-component mode |
4070 |
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in the mode of Ox/Hy Engine |
4510 |
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in the condensate accumulation mode |
4880 |
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Characteristic size of the heat exchanger cells: |
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length, m |
3.5 |
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diameter, mm |
2.0 |
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wall thickness, mm |
0.4 |
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Starting mass, t |
72.3 |
Payload mass, t |
2.5 |
Payload K-t |
0.035 |
Number of launches per year |
9 |
Rocket block of the 1st stage: |
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starting mass, t |
59.03 |
fuel |
kerosene + l. oxygen |
LRE |
RD-120 (forced) |
relative mass |
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construction and compressed gases |
0.083 |
Rocket block of the 2nd stage: |
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starting weight, t |
9.02 |
fuel |
l. hydrogen + l. oxygen |
LRE |
RD-0126 |
relative mass |
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construction and compressed gases |
0.143 |
fuel filling k-t |
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when refueling |
0.93 |
Mass of the correction unit, t |
0.24 |
Launch scheme
The launch of spacecraft to the GSO is carried out in several stages:
1. The launch vehicle, which includes the UUS,
puts the spacecraft into a reference elliptical orbit, at the apogee touching the orbit of the ORC, while
the spacecraft does not separate from the last stage (UUS).
2. With the help of a corrective
The propulsion system bundle of the spacecraft-UUS is transferred to the orbit of the ORC, approaches it and coordinates the mutual
movement.
3. The spacecraft-UUS bundle is brought
close to the ORC at a minimum speed, after which it is captured by the orbital
complex.
4. The spacecraft and UUS are undergoing
maintenance at the ORC, while the spacecraft systems are checked, and the UUS
is refueled. The equipment that is not required for the
subsequent flight is removed from the UUS and the spacecraft, and the most valuable part of it is supposed to be returned to
Land for reuse.
5. The spacecraft-UUS bundle is separated
from the ORC, diverted to a safe distance and, according to a two-pulse scheme, enters a
close to geostationary orbit.
6. The spacecraft is separated from the UUS,
the objects are separated so that the UUS is in an orbit that is safe for
geostationary satellites.
System productivity
The system is focused on close to modern cargo traffic on GSO. The year 2000 was taken as a reference point, in which 9 launches were carried out (all successful). Proton launch vehicle with geostationary satellites. Some of them were put into a geotransfer orbit, where their mass was ~3600 kg. After the additional life on the GSO, their mass decreased to ~2600 kg, but up to 100 kg of them accounted for the excess mass of the propulsion system after the fuel was produced. When directly launched to the GSO, the mass of the satellites was ~2400 kg. Approximately, the Proton in 2000 provided a useful cargo flow on the GSO ~2500×9=22500 êã.
If such a cargo flow passes through the ORC, then, when using single-use carriers with UUS, 67,500 kg of l.oxygen-l.hidrogtn fuel will be required for refueling. This amount of components at their ratio of 6 can be obtained from 86786 kg of water. If the electrolysis-cryogenic unit will work 335 days a year, then at the average annual illumination level kîñâ.ñã=0.65 the fuel production rate will be 3.6 g/s (kîñâ.ñã=Tîñâ/T, ãäå Tîñâ - total time of visibility of the Sun for the total number of orbital periods in a year, T - the sum of the total orbital periods in a year). At the same time, an excess of 19286 kg of oxygen per year is formed. After its use for the needs of the life support system, as well as for maintaining the altitude of the ORC orbit and refueling transport ships, there will be another 12-15 tons, which opens up additional opportunities. For their implementation, additional fuel will be required, delivered by general-purpose cargo ships. The most suitable component is methane, but the use of other fuels or their mixtures with water is also possible.
About the prospects of the proposed launch vthicle
Is the proposed launch vehicle capable of flying, i.e., removing the cargo of the declared mass?
The thrust of the high-altitude RD-120 is 75 t, the starting thrust of the forced RD-120, presumably, will be able to raise
to 95 t. But even this may not be enough to provide an acceptable
thrust-to-weight ratio of the launch vehicle. Then you will need to slightly weigh down the 1st stage, rigidly
install the RD-120 and add a steering LRE, which will provide the missing thrust.
Is the proposed launch vehicle "weak" for the spacecraft in the near future?
Currently, there are two trends in the development of geostationary satellites:
- saturation of opportunities, increase of signal power and number of communication channels;
- miniaturization of the element base of the structure and devices.
These trends are mutually compensated, as a result, a significant increase in capabilities is accompanied by a slight increase in the mass of satellites. But still, the orientation to the prospective mass of satellites of 2.5 tons is wrong. We can expect that a sharp increase in the mass of satellites will not occur, but the capabilities of the proposed launch vehicle will be insufficient. What should I do? The domestic NK-33 LRE closest in power to the RD-120 is not produced, but can be taken from stocks that are limited. Next – RD-191-has twice as much thrust as the maximum forced RD-120. It is not particularly necessary, since there is an alternative – to use two RD-120 (probably not forced). But in any case, this will mean the development of a different launch vehicle and a sharp increase in the mass of the launch vehicle. But, let's say, you need to bring out a satellite with a slightly larger mass than the available launch vehicle is capable of. Then it is proposed to resort to a method tested in other countries – the use of mounted launch boosters. And here it turns out that the cost of forcing the launch vehicle in this way is several times less than the similar costs of direct removal.
This is a characteristic feature of the proposed launch vehicle system. The cost of removing the spacecraft to the GSO is weakly sensitive both to the costs of forcing the launch vehicle, and to the costs caused by its redundancy. Therefore, both ways-the creation of a launch vehicle of the smallest dimension with possible further forcing and the creation of a launch vehicle with redundant capabilities are in the field of competitiveness of the system.
Competitiveness and value of the system
Competition in the field of launch services is specific
and will not determine the future
of the proposed system. Another thing is obvious – the system will first of all
compete with Russian heavy carriers, and already in the second – with
foreign ones. We can only try to compare the theoretical profitability
of the proposed method and direct withdrawal. At the same time, we should focus on
the somewhat distant future, in which we assume the existence
of reusable first stages of the launch vehicle.
Let us consider the comparative characteristics of the direct launch from the equator and two
variants of the launch using the ORC located in orbits with an inclination
of 0° and 51° (see Table). The launch to the reference orbit is carried
out by a two-stage launch vehicle, the first stage of which is reusable, returned to the
launch site, uses dense fuel components and has the specific characteristics
of the Baikal accelerator, the second stage used l.oxygen/l.hidrogen. The scheme of the flight
to the GSO is two–pulse. For the flight from the reference orbit to the ORC, a propulsion system on dense
components is used, which is separated at the station, and direct output to the GSO is carried out
an additional oxygen-hydrogen upper stage (OHUS). The equivalent
cost of removal is calculated on the condition that the specific cost of the refueled
components is less than the specific cost of removing the cargo to the LEO by conventional
means by 2 times for the ORC(51°) and by 2.15 times for the ORC (0°).
|
Inclination, ° |
0 |
0 |
51 |
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Starting mass |
100 |
100 |
100 |
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Weight of the 1st stage |
76.6 |
76.6 |
76.6 |
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incl. used fuel consumption |
66.4 |
66.4 |
66.4 |
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Specific impulse in void, N·s/kg |
3300 |
3300 |
3300 |
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Characteristic speed, m/s |
3600 |
3600 |
3600 |
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Fairing mass |
0.92 |
0.80 |
0.73 |
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Mass of the 2nd stage |
18.4 |
18.5 |
18.9 |
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incl. used fuel consumption |
15.8 |
15.9 |
16.2 |
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Specific impulse, N·s/kg |
4600 |
4600 |
4600 |
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Characteristic velosity, m/s |
5600 |
5600 |
5800 |
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Propulsion system mass |
|
0.52 |
0.50 |
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incl. used fuel consumption |
|
0.39 |
0.38 |
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Specific impulse, N·s/kg |
|
3300 |
3300 |
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Characteristic velosity, m/s |
|
200 |
200 |
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Mass of the 3rd stage |
2.69 |
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incl. used fuel consumption |
2.31 |
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Specific impulse, N·s/kg |
4600 |
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Characteristic velocity, m/s |
3936 |
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Refueling mass |
|
8.1 |
10.9 |
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Specific impulse, N*s/kg |
|
4600 |
4600 |
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Characteristic velocity, m/s |
|
3857 |
4808 |
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K-t filling of tanks to the 2nd inclusion |
|
0.14 |
0.25 |
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Payload mass |
1.32 |
3.51 |
3.21 |
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Equivalent to the cost of delivery 1 kg of payload |
1.28 |
1.00 |
1.35 |
Direct launch from the equator and output via the OS (51°) are approximately equally effective.
The obtained values differ by 7% in favor of the equatorial launch, but it
does not take into account the cost of the OHUS (up to 10% of the cost of the launch vehicle). Cost of propulsion system, dry weight
which, with an equal mass of payload, is an order of magnitude less than that of OHUS, can be ignored.
For direct removal from the equator of the same load will require 2.4 times more
powerful launch vehicle.
The launch using the equatorial OS seems to be the most profitable, but due to the
insufficient value of the characteristic speed of the flight from the equatorial LEO to the GSO get an excess UUS. As a result, it is not
fully refueled, has an excessive weight of the structure and before the last turn-on has a small filling of the tanks – 14%, which can create a problem of reliable
propulsion system start-up. The alternative is a re-sized first stage and reduced payload mass.
However, upon careful consideration of the issue, it becomes obvious that
the operation of the OS in equatorial orbit will require the creation from scratch of the entire
infrastructure to ensure its operation. This is not only the actual
spaceport, but also command and measurement facilities, communication systems,
crew rescue service, and much more. Just look at the map to make sure that,
for example, in the equatorial zone there are no convenient landing places for reusable
vehicles on the flight path. Weather, landscape, and local
socio-political conditions are also generally unfavorable. Total costs
the cost of creating an infrastructure for the Equatorial OS and the cost of its operation
seem so significant that today there is no reason to discuss such
a possibility. The higher efficiency of the equatorial-based RST is also questionable. It is economically most advantageous that the airfield from which
the launch flights are conducted is located in the immediate vicinity of the
EFT production site. For the equator, this means the deployment of high-tech
production, which has never been there (at the equatorial spaceports
only assembly operations are performed). Significant capital
investments will be required, and the costs associated with the isolation of such production from large
industrial centers can devalue the higher load capacity of the RST.
For a fully reusable RST, the location does not play a significant role,
but this is already a different level of cargo flow.
The considered example suggests that for launches on the GSO, when using
the proposed launch vehicle system, the near-equatorial spaceports do not have a decisive
advantage over the mid-latitude ones, and therefore, such a system is especially
relevant for our country. There is a choice: to achieve similar
results, develop the technologies of domestic enterprises or make capital
investments in other territories. This very possibility gives grounds for
Roscosmos does not invest in the launch infrastructure in the near-equatorial
zone.
Launch reliability
There are prerequisites for the fact that
the theoretical reliability of the proposed method of derivation on the GSO will be higher than
with direct derivation. Consider the most likely failure situations and their
consequences. Note that the reliability of the RST practically does not affect the performance
of the task, since it is easy to accumulate water at the ORC in sufficient quantities to
wait out supply interruptions.
Let us consider the consequences of several typical emergency situations for the launch vehicle with UUS and for the launch vehicle of direct excretion on the GSO, discussed above.
|
Situation |
Conventional launch vehicle |
launch vehicle with UUS |
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Failure of the 1st stage enjine at start-up |
Start stop |
Start stop |
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Failure of the 2nd stage enjine |
Loss of cargo |
Loss of cargo |
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Non-achievement ORC by means of UUS |
The stage is absent |
Possible loss cargo delivery |
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Engine failure at start-up with NCO |
Loss of cargo |
Loss of UUS |
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Enjine failure at the second orbital launch |
Loss of cargo |
Loss of cargo |
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Occurrence of problems on board the cargo during removal |
The situation will continue |
Possible repairs |
Statistics show that the most common launch vehicle accidents occur due to engine failure, and engine accidents in most cases occur during transients associated with their start and stop. In the first case, the engine runs for no more than 2 seconds, which makes it possible to stop the start in the event of an accident at the first stage, and in the second , the accident is more often hidden and manifests itself at the next start. Thus, most of the hidden defects appear in the first two launches, and an accident in subsequent launches is less likely. Media intended for direct launch, has one more missile unit, therefore, the probability of having manufacturing defects in it is higher. When the propulsion system is removed, its first and second stages are started once, and the propulsion system of the US is started twice.
In the proposed method, the first-stage propulsion system is started once, and the UUS propulsion system is started three times. At the same time, re-launch The UUS is made during a joint flight with the ORC, and in the event of its failure, the cargo is not likely to be lost.
The proposed method has one additional step - the transportation of a bundle of UUS-payload to the capture zone of the ORC. The experience of operating orbital stations has shown the potentially high reliability of such an operation. But if the ORC is not reached, it will not be difficult to "catch" the bundle, if special means and methods are provided for this purpose.
It should be noted that the power of the engines of the 1st and 2nd stages of the traditional carrier, even at the equatorial start, is ~2.5 times more. This makes it possible to conduct a larger volume of tests and other measures aimed at improving the reliability of the carrier with UUS at the same cost. In reality, however, the development of new media will build on existing technical advances. Therefore , the actual reliability of the carrier will depend on the reliability of the specific engines installed on it. In addition, the comparison is conditional, since the carriers of the considered type are not used for launching vehicles on the GSO. Currently, the typical launch vehicle for direct excretion is two-stage with side accelerators that put the spacecraft into a geotransfer orbit. Due to the differences in circuit solutions and engine types, a correct comparison of the reliability of using these carriers and the proposed method will look more difficult, especially since you will have to compare the hypothetical launch vehicle with the real ones.
Future challenges for the GSO
The prospects of the system will depend not on the cost of withdrawal, but on its ability to develop and solve new problems. The proposed system has considerable potential for development, which can be implemented in an evolutionary way (increasing the power of the ORC, intensifying RST flights, involving more and more powerful launch vehicle in the system). The development of the system may result in the implementation of new opportunities to expand activities in the GSO. The main ones are:
It is likely that the activities under clauses 2 and 3 will never become relevant. But it should be planned for the following reason. This activity means reaching a new, higher level of system relations, and the development of all 5 points – a bid for dominance in the geostationary orbit. The fact that the proposed launch vehicle system can not only subdue the launch services market, but also bring down the production of communication satellites shows how strong the position of this level is. At the same time, the activity of the system will be beyond the influence of artificial factors. To turn such properties of the system to the benefit of the Russian space industry is a matter of technology and the art of foreign economic policy..