The system of means of launching spacecraft to

geostationary orbit


 Composition and parameters

 Launch scheme

 System productivity

 About the prospects of the proposed launch vehicle

 Competitiveness and value of the system

 Launch reliability

 Future challenges for the GSO


Composition:

Orbital refueling complex;

Reusable space tanker;

A launch vehicle that includes a universal

upper stage.

   

Cargo flow on GSO, t/year

22.5


Electrolysis and cryogenic production
complex at the orbital station

   

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


Reusable space tanker

   

Carrier aircraft: special modification of the AN-124 or more

light aircraft.

Rocket block:

0.035

   

Starting weight, t

69.0

Payload mass, t

3.1

Number of launches per year

30

   

Fuel mass for the main propulsion system, t:

 

liquid oxygen

51.2

liquid hydrogen

5.8

керосина

5.2

   

Specific impulse of the propulsion system, N*s/kg:

 

in three-component mode

4070

in the mode of Ox/Hy Engine

4510

in the condensate accumulation mode

4880

   

Characteristic size of the heat exchanger cells:

 

length, m

3.5

diameter, mm

2.0

wall thickness, mm

0.4


Launch vehicle, including UUS

   

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:

 

starting mass, t

59.03

fuel

kerosene + l. oxygen

LRE

RD-120 (forced)

relative mass

 

construction and compressed gases

0.083

   

Rocket block of the 2nd stage:

 

starting weight, t

9.02

fuel

l. hydrogen + l. oxygen

LRE

RD-0126

relative mass

 

construction and compressed gases

0.143

fuel filling k-t

 

when refueling

0.93

   

Mass of the correction unit, t

0.24


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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.

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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 кг.

   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 kосв.сг=0.65 the fuel production rate will be 3.6 g/s (kосв.сг=Tосв/T, где Tосв - 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.

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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 Ц RD-191-has twice as much thrust as the maximum forced RD-120. It is not particularly necessary, since there is an alternative Ц 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 Ц 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.

to the begining 


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 Ц the system will first of all compete with Russian heavy carriers, and already in the second Ц 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 twoЦpulse. 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

Starting mass

100

100

100

Weight of the 1st stage

76.6

76.6

76.6

  incl. used fuel consumption

66.4

66.4

66.4

Specific impulse in void, NЈs/kg

3300

3300

3300

Characteristic speed, m/s

3600

3600

3600

Fairing mass

0.92

0.80

0.73

Mass of the 2nd stage

18.4

18.5

18.9

  incl. used fuel consumption

15.8

15.9

16.2

Specific impulse, NЈs/kg

4600

4600

4600

Characteristic velosity, m/s

5600

5600

5800

Propulsion system mass

 

0.52

0.50

  incl. used fuel consumption

 

0.39

0.38

Specific impulse, NЈs/kg

 

3300

3300

Characteristic velosity, m/s

 

200

200

Mass of the 3rd stage

2.69

 

 

  incl. used fuel consumption

2.31

 

 

Specific impulse, NЈs/kg

4600

 

 

Characteristic velocity, m/s

3936

 

 

Refueling mass

 

8.1

10.9

Specific impulse, N*s/kg

 

4600

4600

Characteristic velocity, m/s

 

3857

4808

K-t filling of tanks to the 2nd inclusion

 

0.14

0.25

Payload mass

1.32

3.51

3.21

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 Ц 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.

to the begining 


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

Failure of the 1st stage enjine at start-up

Start stop

Start stop

Failure of the 2nd stage enjine

Loss of cargo

Loss of cargo

Non-achievement ORC

by means of UUS

The stage is absent

Possible

loss cargo delivery

Engine failure at start-up with NCO

Loss of cargo

Loss of UUS

Enjine failure at the second

orbital launch

Loss of cargo

Loss of cargo

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.

to the begining 


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:


  1. Opportunity a sharp increase in the mass of geostationary satellites. To do this, you will need more powerful UUS and customizing the available media for them. Today, there is no demand for launching such satellites, partly because there is no supply.
  2. Opportunity repair and retrofitting of satellites. With the achieved perfection and reliability satellite equipment cases of its failures are rare, because of this repair work the activity on the GSO itself is irrelevant. From satellite retrofitting the installation of additional solar panels and propulsion systems on them is interesting. Usually, the final service life of the devices is determined by the resource of these devices. the rest of the equipment remains operational. Such operations they may well be conducted without the direct participation of a person. currently, communication satellites are becoming obsolete during their existence morally. But this situation will not be long - lasting, since the limit is the perfection of satellite communication technologies is already visible.
  3. Opportunity refueling of satellites with components. Refueling the satellite with components, used in the propulsion system, it is much cheaper to install a new propulsion system, but the design of the device must be provided with appropriate devices and mechanisms.
  4. Opportunity collection, localization and information from the orbit of space debris. In the present currently, there are more than 700 large objects in the vicinity of the GSO Ц about 200 working apparatuses, "extinct" apparatuses, upper stages, adapters, etc. large debris. Recently, satellites that have exhausted their resource, they are transferred to the burial orbits, but this solves the problem of cleanliness space is only temporary. If we don't want to exclude in the remote in the future, manned flights on the GSO, then today should not be allowed, to have hundreds of unmanageable objects dangling, colliding, and crumbling in the vicinity of such an important area of space. Not very romantic, but useful work on their collection, passivation, compact storage, and possibly de-orbiting, in the future will be in demand.
  5. A human presence may be required to ensure that items 2-4 are met. The use of ORC makes the real orbital station on the near-geostationary orbit. At the same time, two points should be noted. First, only an occasional visit to the GSO will be required to ensure p. 2-4 , and in the long term it may not be necessary at all. Secondly, the situation on the GSO is radiation-hazardous for the crew, it needs to radiation shelter. Fortunately, the GSO launch can be equated with the achievement of a celestial body, since it is already a bit of a stretch. the surrounding area contains more than 1000 tons of matter, enclosed in several layers. hundreds of abandoned objects. This matter can be used, for example, to create radiation protection.

   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 Ц 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..

to the begining