RU2179941C1 - Space rocket system and method for rendering services in launching space vehicles using space rocket system - Google Patents

Abstract

FIELD: space rocket engineering for putting space vehicle into near-earth orbit. SUBSTANCE: system has launch vehicle, space vehicle with adapter, launching pack, launching pad, launcher, power supply system, ground monitoring equipment, set of conveying and servicing facilities, launching way equipment, and space vehicle flight tracking and data processing stations. Nose fairing of launch vehicle, space vehicle, and adapter form self-contained assembly serving as self-contained nose unit with parts for fastening to launch vehicle; launching pack assembled with launch vehicle constitutes transportable rocket module. Launch vehicle has four booster stages and development stage with solid-fuel engine correction unit. Instrument compartment is located in development stage. Method for rendering services in launching space vehicle using this system is based on conveying space vehicle and launch vehicle to launching pad, their joint docking followed by starting launch vehicle with space vehicle. Prior to do so, assigned parameters of space vehicle orbit are compared with orbit parameters ensured by system. If they are within the range of those proposed by system, transportable space rocket module incorporating launching pack and launch vehicle, as well as nose fairing, space vehicle, and adapter are conveyed to launching pad separately. Then self-contained nose unit is assembled by docking adapter, space vehicle, and nose fairing. Space rocket is assembled by joining self-contained nose unit to launch vehicle and rocket is launched. EFFECT: simplified design of space rocket system, trouble-free launching of space vehicle. 20 cl, 25 dwg

Description

 This group of inventions relates to rocket and space technology and is intended for use in launching spacecraft (SC) into low Earth orbit.

 The global market of launch services for launching commercial spacecraft launches has a steady tendency to constant expansion, since launching spacecraft can solve a number of the most important national economic problems, including the creation of a global communication system with the most remote corners of the country, global environmental forecasting situations, exploration of the bowels of the Earth and much more.

 The launch services market offers a wide range of space rockets, including solid fuel.

 Known space rocket complexes (CRK) containing a modular launch vehicle (LV) with lower and upper stages of the installed in a bunch of respectively large and small rocket engines, as well as a head fairing, spacecraft, vehicles delivering spacecraft and launch vehicles, launch pad and launcher (US 5217188 A, 06/08/1993, US 5172875 A, 12/22/1992). When providing services for launching the spacecraft using well-known airspace control systems, the spacecraft and the launch vehicle are transported to the launch pad, their joint docking and subsequent launch of the launch vehicle from the spacecraft from the launch pad are performed.

 The disadvantages of the known KRC are determined by their structural complexity and disunity of the functional components, in connection with which the known methods of providing services for launching the spacecraft are characterized by difficulties in assembling and transporting the components of the KRC. This determines the resultant substantial material costs.

 Closest to the proposed KRC is a complex containing a booster rocket with booster stages having propulsion systems on solid fuel, pyrotechnic devices, an instrument compartment and a head fairing, a spacecraft with an adapter, a transport and launch container (TPK), an engineered launch pad , launcher, power supply system, ground control and launch equipment, a set of transport and technological equipment, launch track equipment and tracking and processing points Records of spacecraft flight information (RU 2025645 C1, 12.30.1994).

 Closest to the proposed method of providing spacecraft launch services using the ASC is a method based on transporting to the prepared launch pad spacecraft and launch vehicles in containers on specialized vehicles, their joint docking and subsequent launch of the launch vehicle from the spacecraft with the launch sites (see ibid.).

 The disadvantages of this complex and the method of providing services for launching a spacecraft are also associated with an unsatisfactory solution to the problems of building a spacecraft, its assembly and transportation.

 The task of the first of the group of inventions is to create a fairly simple in design and transportable CRC, ensuring the reliable launch of the spacecraft with a fully solid rocket launch vehicle weighing up to 100 tons.

 The problem is solved in that in a missile launcher containing a booster rocket with booster stages, which have solid-fuel propulsion systems, pyrotechnic devices, an instrument compartment and a head fairing, a spacecraft with an adapter, a launch vehicle, an engineering launch pad, and a launch pad installation, power supply system, ground control and launch equipment, a set of transport and technological equipment, launch track equipment and tracking and information processing points about flight of the spacecraft — with the head fairing, the spacecraft and the adapter assembly — an independent assembly unit is formed — an autonomous head unit with attachments to the launch vehicle, and the transport and launch container assembly with the launch vehicle — a transportable launch module, while the launch vehicle is made with four booster stages and equipped with a finishing stage with a solid fuel correction engine, the first three booster stages of the launch vehicle are connected in series using the first and second connections itelnyh compartments, respectively, the latter accelerating fourth step performed by the penile scheme propulsion system disposed within the third connecting compartment, lapping stage is connected to the fourth boost stage via a fourth connecting compartment, the instrument compartment disposed in the honing stage.

 Partial essential features of the invention contribute to the solution of the problem.

 Pyrotechnic devices of the launch vehicle are made in the form of squibs and detonating elongated charges, the actuators of which are electrically connected via the cable network to the instrument compartment.

 The first booster stage of the launch vehicle is equipped with a tail compartment made in the form of an aluminum alloy shell reinforced by stringers, and lattice aerodynamic rudders and stabilizers are fixed on the outer surface of the tail compartment, the connecting first and second compartments are made in the form of conical shells supported by spars, stringers, end faces and intermediate frames, while the lining, stringers and intermediate frames are made of aluminum alloys, and the side members and end frames are made of titanium of nano alloys, detonating elongated charges are stacked along the side members and around the perimeter of the end frames, and their actuators-detonators are electrically connected to the instrument compartment, the connecting third compartment is made in the form of finned conical composite panels interconnected by forming spars, and by the end faces frames, while the side members and end frames of the connecting third compartment are made of aluminum and titanium alloys, and detonating elongated charges are laid along the side nerons and along the perimeter of the end frames, and their actuating mechanisms-detonators are electrically connected to the instrument compartment, the connecting fourth compartment is made in the form of a shell reinforced by stringers and end frames, while the shell sheathing, stringers and end frames are made of aluminum alloys, and the rear end frame - with two rows of docking holes located around the perimeters of concentric circles, a third is attached to the docking holes of the outer circle with the front end the main compartment, and the front end housing is attached with the front end housing of the fourth acceleration stage propulsion system, in the finishing stage the correction engine is located in the aggregate compartment, and the fastening points of the finishing stage to the adapter of the autonomous head unit are located on the front end of its platform.

 The transport and launch container is equipped with an aiming system device, the volumetric power structure of the transport and launch container is made in the form of a shell with front and rear covers and is equipped with a cable network and air duct system, the launch vehicle is equipped with supporting-leading belts, and the instrument compartment of the finishing stage is equipped with a porthole for optical communication with the instrument of the aiming system, mounted on the outer surface of the transport and launch container, the onboard cable network of the launch vehicle has a circuit board pumping communications with the cable network of the transport and launch container, with the launcher in contact with its supporting-leading belts, and the launch rocket itself fastened by the tail section to the rear lid of the transport and launch container.

 An antifriction coating is applied on the inner surface of the power shell of the transport launch container.

 The transport and launch container is equipped with a system for maintaining in its internal volume a predetermined heat and humidity regime.

 The internal cavity of the head unit is sealed.

 By docking to a launch vehicle of an autonomous head unit, a space rocket (ILV) is assembled.

 The task of the second of the group of inventions is to increase the efficiency of services for launching spacecraft from anywhere in the world when using the existing infrastructure, both launch pads and flight paths.

 The problem is solved in that in the method of providing services for launching a spacecraft using the ASC, based on transportation to an engineered launch pad of the spacecraft and the launch vehicle, their joint docking and the subsequent launch of the launch vehicle from the spacecraft from the launch pad, are previously compared assigned orbit parameters of the launched spacecraft with orbit parameters provided by the space rocket complex, if the assigned orbit parameters are found and their accuracy is in the range of parameters proposed by the complex are transported separately to the launch pad, a transportable space rocket module consisting of a transport and launch container and a carrier rocket loaded with fuel and pyrotechnics installed therein, as well as a head fairing, a spacecraft and an adapter, assemble an autonomous head unit by joint docking an adapter, a spacecraft and a head fairing, by docking to a launch vehicle of an autonomous head unit, a space rocket is assembled about the destination, raise it to a vertical position and carry out the launch.

 Partial essential features of the invention contribute to the solution of the problem.

 The assembly of the autonomous head unit is carried out in a separate room equipped with assembly means on the launch pad - the “clean room”, after which autonomous checks of the assembled autonomous head unit are carried out and transported to the starting position.

 The docking of the autonomous head unit at the launching position is carried out to the launch vehicle located on the launcher or on the launch stand, after which the final checks of the assembled space rocket are carried out when the space rocket is in a horizontal position.

They provide a spacecraft mass of 100 ... 500 kg, orbital inclination of 52 ... 98 ^o , orbit altitude of 200 ... 1000 km.

 They provide an error in achieving a speed of 2%, inclination of the orbit of 0.1%, orbit altitude of 1%.

 When the assigned parameters of the orbit are found in the range of the provided parameters, the adapter is designed, manufactured and subjected to ground tests for strength and functioning, and if the test data matches the design, they are transported to the launch pad.

 In the room before the assembly of the autonomous head unit, the gas medium is cleaned with a purity of 600 ... 3600 particles per 1 liter with the corresponding particle sizes of 0.4 ... 6 μm at a moisture content of 44 ... 46%.

 When transporting a transportable space rocket module to the launch pad, the specified ranges of temperature and humidity are maintained in the transport and launch container.

In the transport and launch container maintain the temperature in the range of 5 ... 15 ^o C and humidity in the range of 40 ... 50%.

 Transportation to the launch pad of a transportable space rocket module is carried out in an insulated railroad car or on an An-124 airplane or on a sea ship.

1 shows a launch vehicle;
figure 2 - autonomous head unit;
figure 3 - lapping stage;
figure 4 - autonomous head unit docked to the launch vehicle;
in FIG. 5 - callout I in figure 4 - decaying attachment site of the head fairing;
figure 6 - launch vehicle "Start-1" with a docked autonomous unit;
figure 7 - docked modules I and II of the launch vehicle;
on Fig - module III of the launch vehicle;
figure 9 - space rocket module on a transport unit;
figure 10 - space rocket module launch vehicle in a transport and launch container;
figure 11 - space rocket module before installing on a transport unit;
in FIG. 12 - space rocket module on a transport unit - front view, view along arrow A in Fig.9;
in Fig.13 is a fragment of leader II in Fig.11;
in Fig.14 is a fragment of leader III in Fig.11;
on Fig - view along arrow B in Fig;
in Fig.16 is a view along arrow B in Fig.14;
in Fig.17 is a fragment of leader IV in Fig.11;
in Fig.18 - leader IV in Fig.11;
on Fig - section GG on Fig;
in Fig.20 - machine support;
on Fig - ground instrument module;
in Fig.22 - power supply unit;
in Fig.23 - launcher on a mobile launcher - front view;
on Fig - launch vehicle in an upright position on a movable launcher before launch;
25 is a flight diagram of a launch vehicle.

 Almost all niches are occupied in the launch services market, except for the provision of services with a fully solid rocket carrier rocket weighing up to 100 tons, which is used as part of a ground-based transportable space rocket complex.

 The space rocket complex contains a launch vehicle (Fig. 1), with booster stages having solid propellant propulsion systems, pyrotechnic devices, an instrument compartment and a head fairing, a spacecraft with an adapter, a transport and launch container, an engineering launch pad, launcher, power supply system, ground control and launch equipment, a set of transport and technological equipment, launch track equipment and tracking and processing points inf formations about the flight of a spacecraft.

 The head fairing, spacecraft and adapter assembly formed an independent assembly unit - an autonomous head unit (Fig. 2) with attachment points to the launch vehicle, and the transport and launch container assembly with the launch vehicle - a transportable missile module (Fig. 10 and Fig. 11), while the launch vehicle is made with four booster stages and is equipped with a finishing stage with a solid fuel propulsion correction system, the first three booster stages of the booster are connected in series with each other oschi first and second connecting sections respectively, the latter accelerating fourth step performed by the penile scheme propulsion system disposed within the third connecting compartment, lapping stage is connected to the fourth boost stage via a fourth connecting compartment, the instrument compartment disposed in the honing stage.

 When designing a space rocket complex, it is necessary to determine the structure and method of providing launch services, since the choice of the most preferred launch vehicle and the complex as a whole depends on this.

Commercial launch services can be represented in six groups:
- ensuring the required operating conditions of the LV before docking and after docking the spacecraft with the LV;
- conducting telemetry of parameters;
- ensuring safety in preparation for the docking of the LV with the spacecraft;
- ensuring flight safety of the LV with the spacecraft;
- compliance with environmental safety standards;
- providing the required values of the parameters of the launch of the spacecraft into orbit.

The most important services of the first group are measures to ensure the temperature and humidity of the spacecraft before docking and as part of the launch vehicle, as well as the "purity" of the atmosphere surrounding the spacecraft, which is necessary to maintain the optical systems of the spacecraft in working condition. Before docking the spacecraft with the LV, all assembly and installation work with an autonomous head unit (adapter, spacecraft and fairing) should be carried out in a clean room with a very low dust content in each m ^{3 of} this room. After docking the spacecraft with the spacecraft, the requirements for temperature, humidity and air purity are met due to the design of the autonomous head unit.

 In addition to the listed services of the first group, operations may be performed to refuel the spacecraft remote control with liquid components and liquid nitrogen, subject to the necessary safety measures and rules, as well as electrical checks of the spacecraft and its interface with the launch vehicle.

 The implementation of the second group of services is associated with providing telemetric measurements of the parameters of the head unit and spacecraft, transferring them to the ground station during the flight as part of the LV. Processing these measurements allows us to compare the data obtained with the requirements for flight conditions and the separation of the spacecraft from the launch vehicle.

 The third group of services includes a number of constructive, technological and organizational measures implemented by the service provider (part together with the customer-user of the services) in the process of transporting the spacecraft to the spaceport, assembling the headunit and the launch vehicle as a whole and providing safety and security measures for the spacecraft when transporting it as a part of the launch vehicle to the launch pad, during launch preparation and during missile launch.

 Safety services in flight as part of the launch vehicle (fourth group) are provided by the technical solutions adopted by the developer of the launch vehicle, in particular, for acceleration stage propulsion systems, control system, connecting compartments, stage separation devices, fairing separation, etc.

 This also includes taking into account the degree of development of the LV design and complex elements. It is known that, the higher the design maturity, the greater the cost of services and vice versa. It is important to find a rational level of performance at a relatively low cost, using, inter alia, borrowed technical solutions, systems and assemblies.

 The flight path of the LV in the active section up to the separation of the spacecraft from the LV and the fall zone of the fulfilled stages are chosen in such a way as to use a "exclusion corridor" of limited width, since disturbance of the ecological balance of the environment during accidents can have serious consequences.

These services include the following:
- improving the reliability and safety in the operation of propulsion systems of upper stages, control systems, governing bodies, steering drives;
- borrowing previously worked aggregates and systems from other analogues;
- the transition to the creation of launch vehicles only on solid mixed fuels;
- Creation and application of the most advanced, at the same time well-developed schemes and start systems.

 Finally, the last group of services is associated with providing the required parameters and spacecraft launch characteristics.

 A high accuracy of the spacecraft launch parameters should also be ensured (reduction of errors in the launch velocity vector, spread of disturbance parameters during separation of the spacecraft from the LV, etc.).

 We formulate the tactical and technical characteristics of the spacecraft that are necessary and sufficient to provide launch services for launching spacecraft into low Earth orbits.

 The maximum mass of the spacecraft is assumed to be 400 kg, which allows us to put a wide range of spacecraft into orbits.

 The main parameters of the orbit of the spacecraft, determined by its practical purpose, are within wide limits, therefore, the energy capabilities and the functioning scheme of the spacecraft are selected from the condition that the spacecraft can be put into orbit in the altitude range 200 ... 1200 km (depending on the mass of the spacecraft). For most applied spacecraft, based on the conditions for ensuring the necessary lifetime in orbit, taking into account aerodynamic drag, the orbiting altitude is 350 ... 400 km. At the same time, in most cases for applied spacecraft, the required orbit height does not exceed the range of 900 ... 1000 km.

The orbital inclination is determined by the scientific and applied purpose of the spacecraft and is in a wide range associated with the location of the cosmodrome, from the minimum values at equatorial trajectories to polar and solar-synchronous with an inclination angle of 90 ... 98 ^o .

 The accuracy of the implementation of the orbital parameters is determined by the purpose of the spacecraft. High accuracy of removal contributes to fuel economy of the corrective propulsion system (DU) of the final stage and ensures the operational commissioning of the spacecraft. Especially important is the minimization of deduction errors during the spacecraft launch to replace a failed spacecraft during restoration of the composition of the orbital grouping of spacecraft operating jointly in orbit (for example, satellite communications, space navigation, etc.).

Launch "Start-1" (Fig.6) can provide a sufficiently high level of accuracy of deduction for the present time, which corresponds to the following indicators of the maximum deviations of the orbit:
- in height at the point of completion of elimination - ± 1 km;
- in height at the point opposite to the end of the breeding - ± 5 km;
- by the inclination of the orbit - ± 3 arc minutes.

The impact levels on the spacecraft during the separation are determined from the conditions of maximum satisfaction of the needs of a sufficiently high circle of potential customers, since when these requirements are met, in many cases it is decisive for receiving an order for the provision of launch services. So, for example, the dynamic perturbations of the spacecraft parameters during separation should be minimal, since the spacecraft orientation organs, as a rule, are designed for long-term operation with small values of the control forces. Based on the capabilities of the respective systems and spacecraft, the following level of the angular velocity of the spacecraft after its separation was adopted for the Start-1 LV:
- on the channels of pitch and yaw - not more than - 1 ^o / s;
- on the roll channel - not more than - 1 ^o / s.

Protection of the spacecraft from aerodynamic loads and from aerodynamic heating is also an important factor due to the need to exclude the effect of the incoming flow on non-rigid structural elements of the spacecraft, for example, solar panels, antennas, or optical surfaces. Therefore, for the Start-1 launch vehicle, a technical decision was made, according to which the spacecraft should be placed under the cowling, which is discharged at a high-speed head not exceeding several kg / m ² units and an aerodynamic heating level of not more than 1200 kW / m ² .

The composition of the environment around the spacecraft is determined by the features of the spacecraft, while the compositional requirements are different. Therefore, with the aim of expanding the range of launch services provided during the development of the Start-1 launch vehicle, the problem was solved of providing in the space around the spacecraft before launching a high degree of air purity, and in some cases, a neutral gas (nitrogen) medium while maintaining a narrow temperature range (+15 ... + 25 ^o C).

 The probability of contamination of the spacecraft for the Start-1 launch vehicle was assumed to be minimal, since in some cases the spacecraft developer, based on the need to keep solar panels or elements of optical devices in working condition, makes requirements to exclude the effects of solid combustion products on the spacecraft in flight fuel from operating solid propellant rocket engines and products from the operation of pyrotechnic devices. This requirement applies not only to the latching stage remote control and pyrotechnic devices located in the immediate vicinity of the spacecraft, but also to the spent remote control of the separated LV stages.

 According to the results of design studies for the Start-1 LV, a four-stage was selected as having an advantage in terms of payload compared to a three-stage LV. Launch vehicle "Start-1" (figure 1) is made with four booster stages and contains propulsion systems 1, 2, 3 and 4 on solid fuel of the first, second, third and fourth booster stages, respectively. The launch vehicle is fully assembled equipped with pyrotechnic devices - squibs (not shown) and detonating elongated charges, the actuators of which are electrically connected to the control system devices via an onboard cable network. The first three booster stages of the launch vehicle are connected in series with each other using the first 5 and second 6 connecting compartments, respectively, and the last fourth accelerating stage is made according to the penal scheme, the propulsion system 4 of which is located inside the third connecting compartment 7.

 A design feature of the Start-1 launch vehicle is the placement of the fourth stage solid propellant solid propellant rocket engine inside the connection compartment and its fastening only at the front end, which reduced the weight of the engine housing structure and, accordingly, increased the payload (compared to the "bearing" engine housing) by reducing loads and aerodynamic heating exceptions.

 In propulsion systems of booster stages, a solid explosion-proof mixed metal-containing fuel with a high specific impulse is used. The RDTT cases are made of high-strength composite materials, the central nozzles are partially recessed into the combustion chamber, which ensures a reduction in the length of the LV connecting compartments.

 Composite materials were used in the design of the connecting compartments connecting the solid-propellant solid-propellant stages of the upper stages, along with aluminum and titanium alloys. Detonating elongated charges are installed on the connecting compartments, reliably ensuring the breaking of mechanical bonds between the stages of the launch vehicle and the separation of the compartments into parts for their discharge. The connecting first and second compartments are made in the form of conical reinforced shells. The shells, connecting the first and second compartments are reinforced in the form of spars, stringers, end and intermediate frames, while the shell sheathing, stringers and intermediate frames are made of aluminum alloys, spars and end frames are made of titanium alloys, detonating elongated charges along the perimeter of the end frames, and their detonators are electrically connected to the squibs.

 The connecting first and second compartments each contain four spars, evenly distributed around the perimeter of the compartment.

 The connecting third compartment is made in the form of three finned conical composite panels made of carbon fiber, interconnected by forming spars, and along the ends by end frames, their fins are made in the form of a “wafer”, while the spars and end frames are made of aluminum and titanium alloys.

 In the connecting third compartment, detonating elongated charges are stacked along the side members and along the perimeter of the end frames, and their actuators - detonators are electrically connected to the control system devices.

 The fourth connecting compartment of the launch vehicle is made in the form of a shell reinforced by stringers and end frames, while the shell casing, stringers and end frames are made of aluminum alloys.

 The rear end frame of the connecting fourth compartment is made with two rows of connecting holes located around the perimeters of the concentric circles, while the third connecting compartment is attached to the connecting holes of the outer circumference, and the engine housing of the fourth accelerating stage is attached to the connecting holes of the inner circle.

 The first stage of the launch vehicle is equipped with a tail compartment 8, on the outer surface of which lattice aerodynamic rudders and stabilizers 9 are fixed, fixed with the possibility of rotation and fixation in a given position. Four aerodynamic steering wheels and four stabilizers are installed on the tail compartment, evenly and alternating with the rear end of the compartment distributed along the perimeter.

 At the first stage, operating on the atmospheric part of the trajectory, control is provided by the lattice stabilizers and aerodynamic control wheels that are opened at the start. At the second and third stages of the Start-1 launch vehicle, gas injection devices into the supercritical part of the nozzle were used as controls, and a rotary control nozzle at the fourth stage of the launch vehicle.

 To ensure the required accuracy of launching the spacecraft into a given orbit, the launch vehicle is equipped with a lapping stage 10 (Fig. 1 and Fig. 3) with a solid propellant propulsion system, during operation of which kinematic parameters errors accumulated by the launch vehicle are compensated. A lapping propulsion system located inside the aggregate compartment consists of a solid fuel gas generator, a gas duct system, and nozzle blocks operating on a “pushing” circuit that are brought to the outer surface of the aggregate compartment. Flight control is carried out by redistributing the gas flow between the nozzles of the valve devices. The lapping stage 10 (FIG. 1) connected to the fourth booster stage through the fourth connecting compartment 11 (the lapping stage is shown in more detail in FIG. 3).

 The lapping stage (Fig. 3) is equipped with a propulsion correction unit 12 for solid fuel and contains a platform 13, an instrument 14 and an aggregate 15 compartments, while the devices of the control system of the launch vehicle are located in the instrument compartment, the propulsion installation of the correction is located in the aggregate compartment, and the nodes fixing the lapping stage to the autonomous head unit - at the front end of the platform.

 In order to achieve high launch accuracy and the maximum value of the payload on the LV, specially made small-sized control system equipment was used.

 The head fairing, the spacecraft and the adapter assembly formed an independent assembly unit - an autonomous head unit (figure 2).

 The stand-alone head unit (Fig. 2) contains a head fairing 16 with attachment points to the launch vehicle, a spacecraft 17 fixed therein and is equipped with a replaceable adapter 18 with attachment points with a head fairing and spacecraft, respectively.

 Among the existing spacecraft, one can single out a family designed for geological exploration from orbit using highly sensitive optical telescopes. One of the operational requirements for such spacecraft is its content in a high-purity gas environment, for example, in a nitrogen environment, both during ground operation, and after docking the spacecraft to the launch vehicle and throughout the entire atmospheric section of the active section of the trajectory. Therefore, the urgent task is to develop an autonomous head unit, which allows: firstly, to provide autonomous ground operation of the spacecraft in a high-purity gas environment; secondly, to develop assembly technology with a booster rocket to ensure that the spacecraft is in a high-purity gas environment: thirdly, to protect the spacecraft from the atmosphere in the active section of the trajectory (AUT).

 An autonomous head unit docked to the launch vehicle (Fig. 4) comprises a spacecraft 17, a head fairing 16, an attachment unit 19 of the head unit to the launch vehicle 20, an adapter 18, nodes 21 of the spacecraft attachment to the adapter. The mounting unit 19 of the head unit to the launch vehicle is made in the form of a re-connecting ring 23, on one end of which there are holes for docking with the launch vehicle, and the other end face is rigidly bonded to the lateral outer surfaces of the head shell shell frame 22 using disintegrating mounting units 24, made in the form of brackets 25, which are fastened by explosive bolts 26 (figure 5). The self-contained head unit is equipped with an internal cavity sealing assembly made on the basis of deformable and sealing gaskets 27. The sealing assembly is made in the form of a bottom 28 sealed with an adapter.

 The adapter 17 is made in the form of a power shell with front and rear 29 end frames. The adapter is fastened with a rear frame with a re-attachment ring 23 attachment points 30. The attachment points 30 are placed around the perimeter, internal to the holes of the head unit to the launch vehicle. The power shell of the adapter with the front frame is facing the inside of the shell volume of the head fairing.

 The spacecraft is located in the cavity of the head fairing. Mounting units 21 KA are located on the front frame of the power shell of the adapter and are made in the form of pyrotechnic locks. Upon arrival at the spaceport, the adapter, fairing and spacecraft are transferred to the training building. The spacecraft, after a full cycle of its preparation, is installed on the adapter and closed with a head fairing. After that, change the air in the volume under the fairing to nitrogen.

 To ensure certain specific conditions (temperature, humidity and other parameters), an autonomous head unit can be placed in an isothermal container. Using such a container, an autonomous head unit is transported to the spaceport.

 The separation of the spacecraft at the end of the active section of the trajectory is as follows. At the command of the control system, the explosive bolts 26 are triggered, and their destruction occurs. The head fairing is separated from the re-connecting ring 23, for example by mechanical pushers (not shown), then the attachment points 21 KA with the adapter are triggered.

 The design of the autonomous head unit and the technology of its assembly provide protection of the spacecraft in a high-purity gas environment both during ground operation, when docked to a launch vehicle, and on the active part of the trajectory.

 Sometimes a launch vehicle 20 assembled with an autonomous head unit 31 (FIG. 6) is called a space rocket.

 Launch vehicle "Start-1" (Fig. 6) can also be formed of three modules. Module I - 33 (Fig. 7), (the first, second, and third stages of the launch vehicle) can be borrowed, for example, from a previously developed launch vehicle removed from combat duty. Module II - 32 (the fourth stage of the launch vehicle) can be borrowed, for example, from another rocket made only in prototypes. Module III - 34 (Fig. 8) (the finishing stage of the launch vehicle and the self-contained head unit) may contain elements of both a new development and partially borrowed from other missiles, for example, the instrument compartment with control system devices.

 In order to protect against adverse climatic influences and mechanical damage, the launch vehicle is continuously, starting from dispatch from the manufacturer and up to launch, located inside the transport and launch container made of composite materials, forming a transportable space rocket module with TPK (Fig. 9 .. .Fig. 11). Using TPK, the necessary temperature and humidity conditions for the launch vehicle are provided and a mortar launch of the rocket using a pore pressure accumulator is carried out.

 A transportable space rocket module, comprises a transport and launch container (TPK) 35 and a carrier rocket 20 installed in its cavity. The transport and launch container, made in the form of a power cylindrical shell with an onboard cable network (BCS) 36, laid in a protective casing, and air ducts 37 connected to its internal cavity located on its external surface. The internal cavity of the container has an antifriction coating. TPK is equipped with front 38 and rear 39 covers. Instead of the back cover, a bottom 40 with a pallet 41 (Fig. 11) having axial displacement can be installed. On the side surface of the container’s power shell, support belts 42 are fastened rigidly with it with attachment points (not shown) to equipment, such as a transport unit 43 (Fig. 9). In this case, the support belts 42 of the container’s power shell are located opposite the support-leading belts (ORP) 44 of the launch vehicle (Fig. 11). In addition, the launch vehicle is equipped with an onboard cable network 45, closed by a duct, laid along the body of the rocket (along two sides in the yaw plane) and consisting of cables, tear-off and re-connecting connectors. The electrical connection of the onboard control system (SU) of the launch vehicle and the onboard spacecraft system with the TPK onboard cable network and ground control and verification equipment is carried out through the communications undocking mechanism board (RTO) 46 (Fig. 17) located in the aggregate compartment 15 of the finishing stage . This board is connected by fragments of cables from the BCS TPK and BCS booster. The communication of the BCS of the launch vehicle with the BCS of the TPK by means of fragments of cables 47 fastened to the ORP provides for the forced undocking of the BCS of the TPK and the BCS of the launch vehicle during its movement in the TPK.

 A hole 48 is made in the container (Figs. 9, 18 and 19), which is in common optical communication (Fig. 18, Fig. 19) with portholes 49 in the instrument compartment 14 and in the platform 13 and serving for optical communication with the gyrostabilized platform, located in the instrument compartment (not shown), a porthole of which is also located on this optical axis.

 The carrier rocket contacts the inner lateral surface of the TPK cylindrical power shell with its support-leading belts 44 (Fig. 13 ... Fig. 16) made of arcuate annular sections connected by decaying attachment points. In the annular gap between the rear cover of the container and the lower end of the launch vehicle (on the tail compartment), an obturating belt 50 is installed (Fig. 13). It abuts its supporting surface to the inner surface of the TPK and is bonded to the ORP 44 of the first booster stage of the launch vehicle 20 and serves to provide a mortar launch of the launch vehicle from the TPK. The launch vehicle itself is fastened by the lower end of this stage to the TPK back cover (the back cover is not shown) by means of a communication unit made in the form of a terminal ring 51. In this case, the groove of the terminal ring interacts with the flanges of the two rings 52 and 53 in contact with it and with each other, one of which 52 is rigidly bonded to the lower end of the stage of the launch vehicle, and the other 53 by means of pins 54 is rigidly bonded to the back of the container.

 Fixation of the launch vehicle in the TPK in the transverse direction is provided by the AFP, and in the longitudinal direction by the communication center on the tail compartment of the propulsion system of the first booster stage.

 The transport and launch container is equipped with an aiming system device (not shown) mounted on the outer surface of the transport and launch container.

 The Start-1 launch vehicle operating scheme is selected in accordance with the adopted operating scheme for the launch vehicle in the launch vehicle, taking into account the propellant propulsion and the measures to reduce the effects on the spacecraft and improve the launch accuracy, and is characterized by the features listed below.

 All prelaunch operations are carried out when the LV is in a horizontal position. The lifting of the launch container from the LV to a vertical position is carried out at the last minute before launch. The rocket is thrown from the TPK by the pressure of the products of combustion of a special starting powder pressure accumulator. The launch of the first stage solid propellant rocket engine is performed after the launch vehicle completely leaves the TPK.

 All accelerating propulsion systems operate until the fuel burns out completely, while the duration of each remote control is about one minute. After the end of operation, the solid propellant solid-propellant rocket of the first stage of the rocket for 10 ... 20 s flies with an inoperative solid-propellant solid rocket of the second stage.

 Sufficient heights of the spacecraft orbits are achieved by introducing a second, main “pause” of up to 10 minutes (depending on the height of the perigee of the spacecraft’s orbit) before launching the solid rocket engine of the last fourth stage. The necessary angular position and LV turns at the “pause” are ensured by activating the fourth stage of the gas-reactive orientation system (GRS) installed on the solid propellant rocket nozzle. To protect the spacecraft from the force and heat of the incoming stream, the fairing is discharged at the main “pause” section when the altitude is more than 120 km. The last stage solid-state solid propellant rocket is turned on at the end of the main “pause” when it reaches a height close to the perigee of a given orbit. After the main stage of operation of the solid propellant solid propellant rocket engine of the last booster stage is completed, the start-up stage remote control is launched, which also works until the fuel burns out completely. To reduce disturbances, the spacecraft is separated from the launch vehicle with a delay of about 30 s with respect to the moment the remote control operation of the finishing stage is completed. To give the spacecraft the desired orientation by the time of its separation from the launch vehicle at the end of the drop section of the thrust of the advanced remote control, the GRSO is switched on again.

 The main technical solutions for the Start-1 LV correspond to its application as part of a transportable space rocket complex, which practically does not impose restrictions on the choice of a launch site (spaceport).

 All means of a transportable space rocket complex (including LV and TPK) are made in a mobile version and can be transported by sea, rail and air (An-124) vehicles. The most difficult and bulky unit of the complex is a self-propelled launcher. Its mass with the LV in the TPK is 102.5 tons, length - 21 m, width - 3.85 m. When moving, the self-propelled launcher fits into a safe turning radius of 27 m (the width of the corridor when turning is 6 m).

 The launch vehicle arrives at the spaceport in the TPK, which is almost completely assembled and ready for launch. It is only necessary to place a payload on it and make a gas refueling of the gas-reactive orientation system, which is carried out without unloading the LV from the TPK.

 The main work at the spaceport is related to the preparation of the payload (spacecraft) and its installation on the launch vehicle. To carry out these works, the technical and launching positions are equipped at the cosmodrome (it is possible to combine them) with the installation and testing building, launching pad, protected premises of the command post for personnel on launch, and parking lots for the units of the complex.

 For the preparation of modern spacecraft and their assembly as part of the LV head unit, as a rule, equipment in an assembly and test building (or in another structure, the so-called "clean room" with a class of air purity of at least 3600 ... 600 particles per liter of gaseous medium with particle sizes of 0.4 ... 6 microns.

 The launch pad is equipped in engineering and geodesic terms (installation of the foundation of the azimuthal orientation system, three geodetic strong points, divertors, etc.).

 The secure room of the command post is equipped with transmitting and receiving equipment of the system of a single time, a remote control and start-up, equipment of the communication system.

 In addition to the funds placed at the cosmodrome, the launch of the launch vehicle is ensured by the use of telemetry reception tools and information trajectory. With regard to the launch of the Start-1 launch vehicle from the territory of foreign countries, it is possible to receive information using appropriate means deployed at existing and newly organized foreign measuring points.

 If necessary, reception of telemetric information can be carried out by special aircraft - IL-76 aerial measuring points.

 The launch of the spacecraft using the Start-1 launch vehicle became possible due to the fact that simultaneously with the development of the launch vehicle design, the development of ground equipment, which was part of the space rocket, part of which is shown in Fig. 20 ... was carried out. 22 - a support machine 55 (FIG. 20), a ground instrument module 56 (FIG. 21) and a power supply unit 57 (FIG. 22).

The launch of the spacecraft can be carried out having the following material part:
- transportable space rocket module (Fig.9);
- a transport unit with attachment points on it TRKM (not shown);
- launcher stand or movable launcher (not shown);
- launch preparation vehicle (MPP) (not shown);
- universal motor heater (UMP) (not shown);
- a support machine 55 (MO);
- a mobile set of test equipment for the measurement system (PC PASI), which includes two machines of measurement systems based on a three-axle self-propelled chassis each, a diesel-electric unit (ADE) based on a two-axle trailer, a machine for operational processing and transmission of information based on triaxial self-propelled chassis (not shown);
- temporary measuring point (VIP), which includes a machine with equipment for the operational processing and transmission of information on the basis of a triaxial self-propelled chassis and a biaxial trailer with reception and recording equipment:
- ground instrument module 56 (NPM);
- power supply unit 57 (BES);
- tent structure (TKS) (not shown);
- azimuthal orientation equipment (OAO) (not shown);
- spare parts and devices (spare parts) (not shown);
- a set of facilities (CSR) (not shown).

 The ground-based instrument module is intended for electrical checks of the launch pad, NPM and rocket in preparation for launch and for launching the rocket. The ground-based instrument module is a pipe-welded rack that has equipment compartments, four retractable bearings and four swivel wheels, as well as special assemblies for fastening the NPM during transportation at the starting position.

 The NPM equipment includes ground control system equipment, a remote control for the power supply system, a small-sized refrigeration machine, and azimuthal orientation system equipment.

 A transport unit (not shown) is designed to transport the fuel dispenser and ensure the temperature and humidity conditions inside the TPK. The transport unit is a long semi-axle chassis on which two cabs are installed (the left driver's cab is double, the right is single), power hydraulic jacks with base plates, an arrow, compartments with equipment, fuel tanks and other equipment. The transport and handling unit is designed to transport the launch pad and reload the fuel dispenser from the transport unit to the launch pad and vice versa. The transport and handling unit is a two-link road train consisting of a tractor and an active five-axle semi-trailer. The tractor includes a four-axle chassis on which a three-seater cabin, a generator station, a pump station, control equipment, fuel tanks and other equipment are installed. The semitrailer includes a power frame, hydraulic jacks with supports, docking and overload mechanisms, control panels, attachment points of the launch pad to the transport and handling unit and other equipment.

 To ensure power supply to the launch pad and the ground-based instrument module, two power supply units (BES) or two support vehicles (MO) are used.

 The power supply unit is a separate module and is designed to generate and distribute electricity to consumers. The power supply unit includes two power units with a capacity of 30 kW each, and switchgears. The equipment of the BES power supply kit converts a three-phase alternating current 380/220 V to a constant voltage of 28.5 V. The BES power system, consisting of two fuel tanks and pipelines, provides power to diesel units and is located on the same frame as the BES.

 The support machine is intended for power supply of ground equipment units. Power supply to consumers is carried out by a power supply unit mounted on the frame of the Moscow Region. In addition, the Ministry of Defense is designed to provide daily life, recreation, meals for staff in the amount of 10 people, the maintenance and storage of food supplies. All equipment is housed in a body mounted on the same frame as a BES on a four-axle chassis.

The universal motor heater is designed to supply hot air to the heat exchanger, which heats the spacecraft when the rocket is in the TPK without a cover on the launch pad, and heats the air in the container with the rocket when it is found during operation without heating. The air temperature supplied to the heat exchanger from the UMP is maintained at the inlet to the heat exchanger in the temperature range from +50 ^o С to +115 ^o С. The heating capacity is in the range from 35,000 kcal / hour to 180,000 kcal / hour. UMP is mounted on a triaxial chassis. The UMP heater and the heat exchanger are interconnected by air ducts.

 A mobile tent frame structure is intended for placement and maintenance of equipment at the launching position when preparing the rocket and launch equipment for launch. The structure of the TKS includes the following elements: a supporting collapsible frame made of bent steel profiles, a coating of reinforced film material, curtain gates measuring 5 m • 5.4 m (width • height) installed on the ends of the TKS, a sliding mechanism consisting of a rail tracks, track rollers and cables with a winch, electrical equipment for lighting and an electric hoist with a lifting capacity of 2000 kg (power sources are part of ground equipment). TKS is installed on special rail tracks placed on steel sleepers (400 • 100 • 600 cm), on a previously prepared site with a hard surface (15 • 14 m).

 Preparation of the launch pad and equipment for launching the launch vehicle is carried out in the following sequence.

 First, the coordinates of the launch pad are determined, and then the coordinates of the launch position for the subsequent installation of the launch pad with the space-rocket module are determined and the installation locations of the ground equipment elements are outlined.

 Next, the engineering preparation of the launch pad and launch pad is carried out, then a launch pad or a movable launcher and elements of ground equipment are installed on the prepared places and they are electrically and mechanically docked.

 At the starting position, the coordinates of the launch point are determined, designate it and relative to this point determine the line of flight direction and the border of the security zone. A safety zone is an area formed by two semicircles. One semicircle with a radius of 70 ... 100 m centered at the launch point is in the direction opposite to the direction of the flight line. Another semicircle has a radius of 190 ... 210 m and is located in the direction of the flight line.

In the engineering preparation of the starting position, hardening and leveling of its surface is carried out to a value that can withstand pressure within 3.5 ... 4.5 kg / cm ² , and leveling is carried out with a longitudinal slope within 0.9 ... 1 ^o and with a transverse slope in the range of 1.9 - 2.0 ^o .

The maximum specific pressure under the base plates of the launch pad or mobile launcher is 3.5 kg / cm ² , the maximum specific pressure under the bottom of the transport and launch container (during start-up) reaches 3.5 kg / cm ² , the maximum pressure under the base plates of the transport unit - up to 4.0 kg / cm ² , the specific pressure under the supports of the transport and handling unit reaches values up to 4 kg / cm ² .

 The geodetic data at the launching point are determined with the following standard errors: square rectangular coordinates - 20 m, geodetic height - 5 m, gravity acceleration - 2 mGal, vertical line deviation components - 2 arcsec, astronomical azimuths of reference directions - 5 arcsec . Geodetic data are determined in the coordinate system of the common earth ellipsoid adopted for a given country. Coordinates, gravity acceleration, height and components of the vertical line deviation are determined for the launch point, which is the projection of the center of the triaxial gyro-stabilized rocket platform onto the physical surface of the Earth. The indicated point must be marked on the ground.

 After that, in the direction of the flight line of the rocket, at a prepared starting position from the launch point, a tent frame structure is mounted and the rails included in the TCS set are laid along the direction of the flight line of the rocket. The rails have a length of 200 ... 210 m.

 Outside the border of the security zone, geodetic reference points are established.

 The geodetic reference point is a pipe with a plate welded to it with anchors for attaching the SAO support or sighting milestones. The pipe of the geodetic reference point is buried in the ground to the mainland, i.e. 1.5 ... 1.6 m below the depth of maximum freezing (thawing) of the soil.

 The launch pad or movable launcher is oriented with the longitudinal axis along the flight direction line of the rocket and mounted on embedded elements in such a way as to ensure that the launch point is aligned with the projection onto the physical surface of the launch position of the triaxial gyro-stabilized rocket platform when it is on the launcher and after lifting it into vertical position. After that, on different sides of the flight direction line and on different sides of the launch point within the safety borders, rocket protection divertors are installed on the launch pad.

 Launch preparation machine is installed on the launch pad at a distance of 100. ..150 m from the tent frame structure in the direction opposite to the missile flight line. At the same time, a protective hopper is equipped at the installation site of the launch preparation machine.

 Other elements of ground equipment, including a diesel-electric unit, a mobile set of test equipment for the measurement system, a machine with operational processing and information transmission equipment, reception and recording equipment, are installed on the launch pad in such a way that their installation sites are connected to the launch position by road, along which temporary racks are placed for hanging and securing cables.

 The mobile launcher (Fig. 24) is mounted on supporting hydromechanical jacks 58. The transport and launch container 35 with the space rocket located in it has been moved to the vertical position by means of a hydraulic cylinder 59, while the front cover of the TPK has already been reset. The pH is before launch. After lifting the TPK to a vertical position, a command is sent from the control system to activate the powder pressure accumulator located in the rear volume of the transport-launch container. The booster rocket is pushed out under pressure of combustion products from the TPK. Before starting the powder pressure accumulator in the TPK, the mechanical bonds between the tail of the rocket and the transport and launch container are destroyed, and when the launch vehicle moves, its support-leading belts are subsequently reset.

The gas pressure of the powder pressure accumulator in the TPK is in the range of 2.5 ... 4.5 kgf / cm ² . To ensure launches of spacecraft, the previously established infrastructure for providing launches of spacecraft and the flight path of spacecraft is used.

 For the equipment of the route of launches and tracking and processing information about the flight of the spacecraft, newly developed equipment and newly developed flight paths of spacecraft are used.

 For spacecraft launches, a previously equipped engineering platform from the existing infrastructure for providing spacecraft launches can be used.

 To ensure the launches of spacecraft, the newly equipped engineering platforms for launches of spacecraft are used.

 The method of providing services for launching spacecraft using a space rocket complex is based on transporting to the prepared launch pad of the spacecraft and the launch vehicle, jointly docking them and the subsequent launch of the launch vehicle with the spacecraft from the launch pad, while previously assigned orbital parameters of the launched spacecraft with orbital parameters provided by the space rocket complex, if found The assigned orbital parameters and the accuracy of their implementation in the range of parameters proposed by the complex of transported separately to the launch pad transportable space rocket module, consisting of a transport and launch container and a launch vehicle mounted therein, equipped with fuel and pyroelectric means, as well as a head fairing, a spacecraft and adapter, carry out the assembly of an autonomous head unit by jointly docking the adapter, the spacecraft and the head fairing, by docking a space rocket is assembled to the launch vehicle of the autonomous head unit, raised to a vertical position and launched.

 The assembly of the autonomous head unit is carried out in a separate room equipped with assembly means on the launch pad - the "clean room", after which autonomous checks of the assembled autonomous head unit are carried out and transported to the starting position.

 The docking of the autonomous head unit at the launching position is carried out to the launch vehicle located on the launcher or on the launch stand, after which the final checks of the assembled space rocket are carried out when the space rocket is in a horizontal position.

 If the assigned orbit parameters are within the range of parameters proposed by the system, the adapter is designed, manufactured and subjected to ground tests for strength and functioning, and if the test data matches the design, they are transported to the prepared engineering site.

 25 is a flight diagram. According to this scheme, space rockets leave their TPK under the action of the gas pressure forces of the powder pressure accumulator (mortar launch) 60, after which the first stage propulsion system is launched. After the fuel burns out, the ДУ1 is separated, the ДУ2 is launched and С01 is separated. After the fuel burns out in the DU2, the DU2 and CO2 are separated and so on until the KA61 is separated.

The range of orbit parameters is - the displayed mass of the spacecraft is 100 ... 500 kg, the inclination of the orbit is 52 ... 98 ^o , the height of the orbit is 200 ... 1000 km.

 The error in the implementation of the range of orbit parameters proposed by the system is - 2% speed, 0.1% orbit inclination, 1% orbit height.

 In the room before the assembly of the autonomous head unit, the gas medium is cleaned with a purity of 600 ... 3600 particles per 1 liter with the corresponding particle sizes of 0.4 ... 6 μm at a moisture content of 44 ... 46%.

When transporting to a site prepared in engineering terms for a transportable space rocket module in a transport launch container, the specified ranges of temperature and humidity are maintained: temperature in the range of 5 ... 15 ^o C and humidity in the range of 40 ... 50%.

 Transportation to a site prepared in an engineering sense for a transportable space rocket module is carried out in a railway insulated car or on an An-124 airplane or on a sea ship.

Claims (20)

 1. A space rocket complex containing a launch vehicle with booster stages having solid-fuel propulsion systems, pyrotechnic devices, an instrument compartment and a head fairing, a spacecraft with an adapter, a launch vehicle, an engineered launch pad, a launcher, power supply system, ground control and launch equipment, a set of transport and technological equipment, launch track equipment and tracking and information processing points mission on the flight of the spacecraft, characterized in that the head fairing, spacecraft and adapter assembly formed an independent assembly unit - an autonomous head unit with attachment points to the launch vehicle, and the transport and launch container assembly with the launch vehicle - a transportable space rocket module, while the launch vehicle is made with four booster stages and equipped with a finishing stage with a solid fuel propulsion correction system, the first three booster stages are no The plugs are connected in series with each other using the first and second connecting compartments, respectively, the last fourth booster stage is made according to the penal scheme with a propulsion system located inside the third connecting compartment, the finishing stage is connected to the fourth accelerating stage through the fourth connecting compartment, the instrument compartment is placed in the finishing stage .  2. The space rocket complex according to claim 1, characterized in that the pyrotechnic devices of the launch vehicle are made in the form of pyro-cartridges and detonating elongated charges, the actuators of which are electrically connected via the cable network to the instrument compartment.  3. The space rocket complex according to claim 1, characterized in that the first booster stage of the carrier rocket is equipped with a tail compartment made in the form of an aluminum alloy shell reinforced by stringers, and lattice aerodynamic rudders and stabilizers are fixed on the outer surface of the tail compartment, connecting the first and second compartments are made in the form of conical shells, supported by spars, stringers, end and intermediate frames, while the casing, stringers and intermediate frames are made of aluminum alloys, and side members and end frames made of titanium alloys, detonating elongated charges are laid along the side members and along the perimeter of end frames, and their actuating mechanisms-detonators are electrically connected to the instrument compartment, the connecting third compartment is made in the form of ribbed conical composite panels, along the forming spars, and along the ends - by the end frames, while the spars and end frames of the connecting third compartment are made of aluminum and titanium alloys, and detonating elongated charges are laid along the side members and around the perimeter of the end frames, and their actuators-detonators are electrically connected to the instrument compartment, the connecting fourth compartment is made in the form of a shell reinforced by stringers and end frames, while the shell sheathing, stringers and end frames from aluminum alloys, and the rear end frame - with two rows of docking holes located around the perimeters of concentric circles, to the docking holes outside the third circumferential compartment is attached with the front end to the front circumference, and the front end housing of the fourth acceleration stage propulsion system is attached to the connecting holes of the inner circumference, the correction motor installation is located in the aggregate compartment, and the fastening units of the final stage are attached to the adapter of the autonomous head unit at the front end her platform.  4. The space rocket complex according to claim 1, characterized in that the transport and launch container is equipped with an aiming system device, the volumetric power structure of the transport and launch container is made in the form of a shell with front and rear covers and is equipped with a cable network and duct system, a launch vehicle equipped with supporting-leading belts, and the instrument compartment of the lapping stage - a porthole for optical communication with the aiming system device mounted on the outer surface of the transport and launch container, onboard abelnaya booster charge network has disconnection mechanical communication with the cable network transport and launch container with an inner side surface which is contacted booster their musculoskeletal leading belts, and itself booster bonded tail compartment with its back cover transport and launch container.  5. The space rocket complex according to claim 1, characterized in that an antifriction coating is applied to the inner surface of the power shell of the launch vehicle.  6. The space rocket complex according to claim 1, characterized in that the transport and launch container is equipped with a system for maintaining in its internal volume a predetermined heat and humidity regime.  7. The space rocket complex according to claim 1, characterized in that the internal cavity of the head unit is sealed.  8. The space rocket complex according to claim 1, characterized in that the space rocket is assembled by docking to the launch vehicle of an autonomous head unit.  9. A method of providing services for launching spacecraft using a space rocket complex, based on transportation to an engineered launch pad of the spacecraft and the launch vehicle, their joint docking and the subsequent launch of the launch vehicle with the spacecraft from the launch pad, characterized in that preliminarily compare the assigned parameters of the orbit of the launched spacecraft with the parameters of the orbits provided by the space rocket complex , if the assigned orbit parameters are found and the accuracy of their implementation in the range of the parameters proposed by the complex is transported separately to the launch pad, a transportable space rocket module consisting of a transport and launch container and a launch vehicle mounted therein, equipped with fuel and pyro-means, as well as a head fairing, a spacecraft and an adapter, assemble an autonomous head unit by jointly docking an adapter, a spacecraft and a head fairing, by docking to a launch vehicle of an autonomous head unit, a space rocket is assembled, raised to a vertical position and launched.  10. The method according to p. 9, characterized in that the assembly of the stand-alone head unit is carried out in a separate room equipped with assembly means on the launch pad - "clean room", after which stand-alone checks of the assembled stand-alone head unit are carried out and transported to the starting position.  11. The method according to p. 9 and 10, characterized in that the docking of the autonomous head unit at the launch position is carried out to the launch vehicle located on the launcher or on the launch stand, after which final checks of the assembled space rocket when the space rocket is located in horizontal position. 12. The method according to p. 9, characterized in that they provide the output mass of the spacecraft 100-500 kg, the inclination of the orbit 52-98 ^o , the height of the orbit 200-1000 km.  13. The method according to p. 9, characterized in that they provide an error in achieving a speed of 2%, orbital inclination of 0.1%, orbit height of 1%.  14. The method according to p. 9, characterized in that when the assigned orbit parameters are in the range of the provided parameters, the adapter is designed, manufactured and subjected to ground tests for strength and functioning, and if the test data matches the design, they are transported to the launch pad.  15. The method according to p. 9, characterized in that in the room before assembling the autonomous head unit, the gas medium is cleaned with a purity of 600-3600 particles per 1 liter with the corresponding particle sizes of 0.4-6 microns with a moisture content of 44-46%.  16. The method according to p. 9, characterized in that during transportation to the launch pad of a transportable space rocket module in a transport and launch container support the specified temperature and humidity ranges. 17. The method according to p. 14, characterized in that in the transport and launch container maintain a temperature in the range of 5-15 ^o C and humidity in the range of 40-50%.  18. The method according to p. 14 or 15, characterized in that the transportation to the launch pad of a transportable space rocket module is carried out in a railway insulated car.  19. The method according to p. 14 or 15, characterized in that the transportation to the launch pad of a transportable space rocket module is carried out on an An-124 aircraft.  20. The method according to p. 14 or 15, characterized in that the transportation to the launch pad of a transportable space rocket module is carried out on a sea ship.

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