Method of launching multi-stage space launch vehicle by means of carrier-aircraft and multi-stage space launch vehicle for realization of this method

FIELD: space engineering.

SUBSTANCE: proposed method includes transportation of space launch vehicle to launching position, preparation for launch, raising the space launch vehicle to preset altitude by carrier-aircraft, separation from carrier-aircraft, stabilization of space launch vehicle and starting the engine plant of first boost stage. Space launch vehicle is transported to launching position in transportation-and-operation container. Then, container is transferred by means of crane to erection trolley, detachable compartments are dismantled and space launch vehicle is transported to carrier-aircraft. Space launch vehicle is secured to carrier-aircraft by means of locks of carrier-aircraft. Space launch vehicle is equipped with boost stages with solid-propellant engine plants, stabilization unit and units for attachment of launch vehicle to carrier-aircraft. It is also equipped with separable tail fairing and lattice stabilizers made in form of cylindrical panels which are secured on it. After bringing the space launch vehicle to preset altitude, locks of carrier-aircraft are opened by command and lattice stabilizers of tail fairing are opened simultaneously. After preset pause, before separation of space launch vehicle, tail fairing with lattice stabilizers is separated from space launch vehicle. Proposed method makes it possible to reduce launch mass and ensure stabilization on flight leg of safe distance from carrier-aircraft till moment of start of 1st stage engine plant.

EFFECT: extended field of application.

7 cl, 5 dwg

 

This invention relates to space technology and can be used in the development of technology start multi-stage space launch vehicle and in the design of multi-stage space launch vehicle.

Currently leading space powers and developing countries are increasing their efforts in the field of exploration and use of outer space using small spacecraft (payload).

The demand for small satellites (MCA) involves the volume of orders on their launch and expansion of this sector of the world market.

Market needs are satisfied by the launch of MCA from stationary spaceport LV medium or heavy class as an additional payload, while a necessary condition is the existence of the order for the removal of large main payload.

The disadvantage of this method start the ICA is a limit on the number of launches due to availability of orders on the excretion of large main payloads.

In practice, currently, the task launch ICA not only from fixed launch sites, but also with the country of the customer or the owner of KA.

Infrastructure analysis of space-rocket complexes (RSC), both stationary and mobile, dormancy is shows to ensure the withdrawal of payload into near-earth space the main parameters of the complex depends on the output payload into near-earth space, the launch method and design space of multi-stage booster. With all of this in the composition of the RCM includes: radio AIDS to receive external information and formation flight, launch vehicle, ground support; system-based spaceport, airfield, launch pad, trained in engineering terms, submarine); means of transportation booster; measuring points of reception of telemetry information about the parameters of the trajectory of the launch vehicle, separation of the levels and branches of the MCA.

The task of the conclusion of satellites in near-earth space using ALAC is met if the MCA was separated from the launch vehicle at a given altitude orbit with an accuracy specified by the terms of the contract. Further maintenance of the ICA provides services that are not members of the ALAC.

Analysis of ballistic schema launch of satellites in near-earth space using a rocket with a motor installations of solid fuel shows that the maximum payload weight for a given height of the orbit can be deduced using ballistic "pause". PR is the branch of the MCA from the booster ideally should take place at a distance of about provide motion along the orbit of the ICA. This circumstance makes it difficult to withdraw the ICA in the near-earth space with the customer premises, as required to place the members of the ALAC (measuring points) outside the territory of the customer.

To solve the problem at the conclusion of satellites in near-earth space with the customer's site using ALAC, which is an Autonomous (independent third parties) and mobile, which requires to solve a number of interrelated problems, which include:

I. the development of the way of the conclusion of satellites in near-earth space and the development of RCM.

II. Development of method startup booster system-based and development booster.

III. Development of a multi-stage space launch vehicle.

Considering the first problem showed that for goal-based selection of relevant motion parameters of the booster and parameters of LV, it is necessary to optimize the composition of the RCM, to solve the problem by creating Autonomous aviation and rocket-space complex (ARCC) and to launch the ICA with the customer's territory. The solution to this problem is devoted to an application entitled "Method output payload into near-earth space using airborne missile and space complex and airborne missile and space complex " on the issuance of p the awning to the invention, served jointly and simultaneously with the present application.

The consideration of the second issue showed that for the purpose you want to take advantage of the aircraft carrier with the technical characteristics of the fighter-interceptor of the fourth generation type MiG-31 at startup booster. The solution to this problem is the subject of this application for the grant of a patent for an invention.

The solution of the third problem was shown that for this goal it is necessary to use multi-stage solid-fuel rocket, since in this case there is no need of filling its components liquid fuel and thereby simplifies ground equipment (not complex for filling), and can also be provided with factory readiness of the rocket and, as a result, high reliability, security, convenience and ease of operation.

Solving the third problem is devoted to an application entitled "Multi-stage space launch vehicle" for the grant of a patent for an invention, which is given jointly and simultaneously with the present application.

The most important phase of RCM is to launch with CH. At this energy the ability of SCC to orbit SPACECRAFT in low-earth orbit largely depend on how fully will be able to use the initial velocity and height, ensure chiweenie aircraft carrier at the time of separation from him rocket.

As considered similar circuit with a load in the cargo compartment of the aircraft KRN, lifting her plane at a predetermined height, the release through the rear hatch of SCC, the descent of SCC on the parachute with its subsequent shooting before starting the propulsion system of the 1st stage.

The disadvantage of this method run SCC is a total loss of kinetic energy, the reported SCC, and at the moment when the propulsion system of the 1st stage of SCC acquires a negative vertical speed, for compensation which takes some energy to SCC after the start of the propulsion system of the 1st stage. Thus, this scheme is used, and it is not full, only the height gain, reported the rocket plane.

The closest to the proposed method output ľa in low earth orbit is the way with land planning KRN before starting the propulsion system of the 1st stage of SCC involving the use of the wing to maintain the required direction of the start of the trajectory of the rocket. This scheme is implemented in the rocket-space complex with aviation start

- American project "Pegasus" when using non-maneuverable subsonic aircraft carrier.

The disadvantage of this method run SCC is significantly heavier construction of the 1st stage in the mouth the wing unit and difficulties associated with the arrangement of such a wing.

The objective of the first group of inventions is the expansion of the scope of SCC, the way its run, while ensuring stabilization of the flight safe removal from the aircraft carrier to launch propulsion 1st step.

The problem is solved in that in the method of starting a multi-stage space launch vehicle using aircraft carrier, which consists in the transportation of SCC on the starting position of the rocket, the preparation of SCC to launch, the rise of SCC at a predetermined height of the aircraft carrier, the Department KRN from the carrier aircraft, stabilization of SCC and the start of the propulsion system of the first booster, exercise equipment KRN detachable tail fairing and mounted thereon folded lattice stabilizers and transport Assembly to the starting position using the transport operating container, overload valve operating freight container with SCC on the transport Assembly cart, take off the detachable compartments operating freight container and transported KRN to the plane of the carrier, shall mount KRN to the plane of the carrier to lock the carrier aircraft, and after lifting KRN aircraft carrier at a predetermined height on the disclosure castles plane is the La simultaneously open lattice stabilizers of the tail fairing and after checkout pause after separation KRN from the aircraft carrier is separated tail fairing lattice stabilizers from SCC.

The task promote private significant features of the invention.

The rise of SCC carried out in a supersonic high-altitude fighter-interceptor fourth generation MiG-31.

Separate KRN from the carrier aircraft in the pitching plane at a height of 15...19 km, while retaining the speed of the aircraft...630 550 m/s and angle of cabrerana 15...35°.

Start the propulsion system of the first step carried out after a calculated pause 5...6 after separation KRN from the carrier aircraft.

The technical task of the second group of inventions in the structure of SCC is to expand the scope of its application by reducing launch mass and ensuring the stability of the flight safe removal from the aircraft carrier to launch propulsion 1st step.

The technical problem is solved by the fact that multi-stage space launch vehicle, containing the upper stages are connected serially via a connecting compartments and equipped with propulsion units for solid fuel, as well as the stabilization device and the attachment points to the plane of the carrier, characterized in that it is equipped with detachable tail fairing and mounted thereon lattice stabilizers, made in the form of a cylindrical panel, with hosto the second fairing performed with a truncated conical shell, supported by end frames, from the smaller base of the conical shell is closed by a spherical segment, on the larger end frame posted by disintegrating the attachment points of the tail fairing to the free end of the first booster, and on a smaller face frame installed nodes fixation, turning and fixing lattice stabilizers, on the lateral surface of the conical shell includes a cavity that matches the shape and dimensions of the lattice stabilizers, in the folded and locked condition, each lattice stabilizer is placed in its cavity and drowned.

The task promote private significant features of the invention.

Angle polestar conical shell of the tail fairing is 15...20°.

The radius of the spherical segment of the tail fairing is 200...300 mm

The effective area of lattice stabilizers is 2.5...3.5 m2.

The inventive method of starting a multi-stage SCC and the inventive multi-stage SCC of the United General inventive concept, the inventive method launch multi-KRN, thus reducing launch mass of SCC and stabilization of SCC on the flight safe removal from the aircraft carrier to launch the motor in the setting of 1-St stage.

A specific example of the method of starting and design of SCC is illustrated in figure 1-5.

Figure 1 shows operating freight container for the rocket.

Figure 2 shows operating freight container Assembly with the booster without removable compartments operating freight container.

Figure 3 shows the aircraft-carrier Assembly with SCC.

Figure 4 shows a multi-stage SCC discharged from the tail fairing and installed lattice stabilizers in the folded position.

Figure 5 shows a multi-stage SCC discharged from the tail fairing and installed lattice stabilizers in the open position.

Using ground-based technological equipment booster loaded with fuel and procreative, operating freight container, 1, is transported from the manufacturer to the base airfield.

Operating freight container 1, made in the form of pencil case has a cylindrical shape and is divided in two transverse planes of the connector 1 front 2, middle 3 and rear 4 compartments. Front and rear compartments made in the form of a cylindrical Cup with a flat bottom and separated by a plane 5 straight connector with attachment points (the attachment points on the figures not shown)on the upper and lower halves, the middle compartment is also divided by a plane 5 straight connector with attachment points (the attachment points on the figures not shown) on the base 6 and the top panel 7.

The mechanical connection of the missile from the aircraft carrier is realized by means of two belts fasteners, each of which is made in the form of a bandage comprising two half-rings (figures not shown). In the joints of semirings posted peretoskali, through which half of the zones rastalkivaya when they reset after separation KRN from the aircraft. Front belt attachment is equipped with a removable power frames. Power frame connected with the grippers on the plane. During ground operation for placing missiles in operating freight container power frame removed (figures not shown).

Kit mechanical connection (figures not shown) provides the laying of SCC on the cradle operating freight container, consolidation of SCC in the container against axial movement and twisting during transportation, holding the shop handling operations with SCC, pinning KRN to the suspension of the aircraft and the Department of mechanical connection elements from SCC after undocking from the plane.

At the base airfield operating freight container with booster set and fixed on the transport and mounting trolley (figures not until the Ana). Transport and mounting trolley KRN transported within the technical zone of the airfield and perform work prior to the suspension of SCC under the aircraft-carrier. Transportation transport and mounting trolley with SCC under the aircraft produced after the dismantling of the removable elements of the transport operation container: front and rear compartments, the top panel of the middle compartment, while SCC is fixed on the base 6, figure 2, middle compartment.

Under suspension of SCC to the plane of the transport and mounting trolley posted on jacks (figures not shown). After the suspension of SCC to the plane checks the reliability of the lockout, KRN unsnaps from the tabs transport and installation of truck and bogie frame is lowered down to ensure the posting of SCC on the locks of the plane (figures not shown).

Aircraft-carrier 8 with SCC 9 shown in figure 3.

To reduce the aerodynamic resistance of SCC in suspension under the plane SCC contains tail fairing 10, figure 4, with folded lattice stabilizers 11.

To stabilize the movement of SCC after separation from the carrier aircraft lattice stabilizers 11, 5, forcibly disclosed.

After lifting KRN aircraft carrier at a predetermined height on the disclosure locks the carrier aircraft (figures not shown) at the same time, what about the open lattice stabilizers of the tail fairing and after check-out pause before starting the propulsion system of the 1st booster KRN separate the tail fairing lattice stabilizers from SCC.

Circuit-construction solutions for the separation system KRN from the carrier aircraft and the launch site of SCC is selected based on the need for inclusion of the propulsion system of the 1st stage of SCC at a distance of more than 60 m from the aircraft to ensure its security through ˜5 after separation of the SCC). Static stability of SCC unmanaged flight after separation from the aircraft to enable propulsion of the 1st stage is provided by the disclosed lattice stabilizers located on the tail fairing KRN discharged before turn on her engine.

Conditions of separation of SCC from CH:

- height - 15...19 km;

- the speed of the aircraft - 550...630 m/s;

transverse overloading of the aircraft at the moment of separation to +2.4...+2,5.

Significant transverse overloading of the aircraft allows the use of special devices to provide a shockless separation of SCC. The disclosure lattice stabilizers served upon rupture of the mechanical connection of SCC with the plane.

In these conditions, after 0.3 s after separation KRN from the aircraft carrier the distance between them exceeds 1 m, allowing this time to complete the disclosure lattice stabilizers. After disclosure of the stabilizers of SCC has a positive margin of static stability and makes slabotermalnye oscillations about the center of mass. At the minimum acceptable margin of static stability 7-8% of the effective area of lattice stabilizers will be 2.8 - 3.0 m2. After separation from the aircraft's on-Board control system KRN current motion parameters based on the conditions of departure from the aircraft at the desired distance and the possibility of subsequent flight control, determines when to reset the tail fairing and run the propulsion system of the 1st stage of SCC.

After separation of the tail fairing from KRN run.

A specific example of execution of a multi-stage space launch vehicle illustrated in figure 1-5.

To implement the method run multi-KRN using the carrier aircraft developed space launch vehicle, shown in figure 4.

Based on the high values required final speed (-8000 m/s)required for SCC, taking into account the fact that part of this speed is reported before the start of the rocket plane for SCC selected scheme with three lively marching steps and apogee stage 12, figure 4, with the propulsion system on solid fuel 13.

To reduce the aerodynamic drag of the rocket on the suspension under the plane and give the rocket static stability at the site of uncontrolled movement from the moment of its separation from the aircraft to start the engine when you remove safely on the distance from the aircraft, the missile contains a tail fairing 10, figure 4, with folded lattice stabilizers 11, forcibly disclosed after reset KRN from the plane. Tail fairing is separated immediately before the start of the propulsion system of the 1st stage of SCC (figures not shown).

Tail fairing is made in the form of a truncated conical shell, supported by end frames. From the smaller base of the conical shell is closed by a spherical segment 14, figure 4, on the larger end frame posted by disintegrating nodes (figures not shown) securing the tail fairing 10 to the free end of the first booster. On a smaller face frame installed nodes fixation, rotation, and mounting (figures not shown) lattice stabilizer 11. On the lateral surface of the conical shell includes a cavity 15, 5, matching the shape and dimensions of the lattice stabilizer 11. In the folded and locked condition, each lattice stabilizer is placed in its cavity and drowned.

Thus, it is shown that the proposed method of starting a multi-stage SCC using the carrier aircraft and the design of multi-stage SCC reduce launch mass of SCC and to take advantage of the carrier aircraft with technical features and advantages of the kami of the fighter-interceptor of the fourth generation type MiG-31 at startup booster.

1. Way to run a multi-stage space launch vehicle (SCC) using aircraft carrier, which consists in the transportation of SCC on the starting position of the rocket, the preparation of SCC to launch, the rise of SCC at a predetermined height of the aircraft carrier, the Department KRN from the carrier aircraft, stabilization of SCC and the start of the propulsion system of the first booster, characterized in that the SCC equip detachable tail fairing and mounted thereon folded lattice stabilizers and transport Assembly to the starting position using the transport operating container, overload valve operating freight container with SCC on the transport Assembly cart, take off the detachable compartments operating freight container and transported KRN to the plane of the carrier, shall mount KRN to the plane of the carrier to lock the carrier aircraft, and after lifting KRN aircraft carrier at a predetermined height on the disclosure locks the carrier aircraft simultaneously open lattice stabilizers of the tail fairing and after check-out pause before separating the SCC from the aircraft carrier is separated tail fairing lattice stabilizers from SCC.

2. The method according to claim 1, characterized in that the starting of the propulsion system of the first stage are carried out by IP is icenii calculated pause 5-6 s after separation KRN from the carrier aircraft.

3. The method according to claim 1, characterized in that the separate KRN from the carrier aircraft in the pitching plane at a height of 15 to 19 km, while retaining the speed of the aircraft 550-630 m/s and angle of cabrerana 20-40°.

4. Multi-stage space launch vehicle, containing the upper stages are connected serially via a connecting compartments and equipped with propulsion units for solid fuel, as well as the stabilization device and the attachment points to the plane of the carrier, characterized in that it is equipped with detachable tail fairing and mounted thereon lattice stabilizers, made in the form of a cylindrical panel, with the tail fairing is made with a truncated conical shell, supported by end frames, from the smaller base of the conical shell is closed by a spherical segment, on the larger end frame posted by disintegrating the attachment points of the tail fairing to the free end of the first booster, and a smaller face frame installed nodes fixation, turning and fixing lattice stabilizers, on the lateral surface of the conical shell includes a cavity that matches the shape and dimensions of the lattice stabilizers, in the folded and locked condition, each lattice stabilizer is placed in its cavity, and with the linen.

5. The booster according to claim 4, characterized in that the angle of polestar conical shell of the tail fairing is 15-20.

6. The booster according to claim 4, characterized in that the radius of the spherical segment of the tail fairing is 200-300 mm

7. The booster according to claim 4, characterized in that the effective area of lattice stabilizers is 2,5-3,5 m



 

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EFFECT: enhanced stability and improved vibration insulation of the apparatus.

7 cl, 8 dwg

Settlement in space // 2264330

FIELD: construction of large-sized structures in space; space engineering.

SUBSTANCE: proposed settlement includes production, living and auxiliary rooms built from lightened modules which are combined in single complex of cylindrical shape with tunnel located along its main longitudinal axis. Said tunnel is embraced by three bodies: main body, body of communication chambers and body of transfer chamber. Pressurized passages are provided between these chambers. Gravitational drive mounted on tunnel is used for rotating the complex in order to form artificial gravity in all rooms. Facing secured on outer surface of guards is used for protection against adverse effect of space. Guards of main body are made from torous members assembled from enlarged space building modules. Main body may be provided with hollow longitudinal and radial stiffening members fastened together and secured to said torous members and to tunnel.

EFFECT: reduced labor consumption and time required for assembly of space structure.

2 cl, 8 dwg

Settlement in space // 2264330

FIELD: construction of large-sized structures in space; space engineering.

SUBSTANCE: proposed settlement includes production, living and auxiliary rooms built from lightened modules which are combined in single complex of cylindrical shape with tunnel located along its main longitudinal axis. Said tunnel is embraced by three bodies: main body, body of communication chambers and body of transfer chamber. Pressurized passages are provided between these chambers. Gravitational drive mounted on tunnel is used for rotating the complex in order to form artificial gravity in all rooms. Facing secured on outer surface of guards is used for protection against adverse effect of space. Guards of main body are made from torous members assembled from enlarged space building modules. Main body may be provided with hollow longitudinal and radial stiffening members fastened together and secured to said torous members and to tunnel.

EFFECT: reduced labor consumption and time required for assembly of space structure.

2 cl, 8 dwg

FIELD: spacecraft temperature control systems.

SUBSTANCE: proposed method includes measurement of temperature of spacecraft structural members and onboard equipment and components of rocket propellant, heating them by celestial body heat and conversion of electrical energy into thermal energy as measured temperatures reach low limits of thermostatting range. In flight, intervals of thermal energy accumulation in propellant components (at excess of thermal energy and electric power on board) and intervals of its free liberation are determined. In case expected magnitude of accumulated energy during predetermined interval exceeds upper level for preset volume of propellant, heat of celestial bodies is accumulated till the end of this interval. Otherwise, excess of electric power generated on board is converted into heat which is delivered to propellant components. In predicting release of thermal energy from propellant components, its residual amount required for maintaining the propellant component temperature within required ranges is determined; temperature of structural members and onboard equipment is also measured. In case this temperature exceeds permissible levels, delivery of heat is discontinued. When temperature of propellant component gets beyond threshold magnitudes, removal of heat from propellant components is discontinued. Otherwise, delivery of heat to thermostattable elements and onboard equipment and/or to points of accumulation of heat for subsequent useful conversion is continued till beginning of next interval of accumulation of thermal energy. Then, thermal energy accumulation cycle is repeated.

EFFECT: enhanced efficiency of accumulation and release of thermal energy; reduced mass and overall dimensions; enhanced heat removal.

5 dwg

FIELD: spacecraft temperature control systems.

SUBSTANCE: proposed method includes measurement of temperature in areas of radiation surfaces of temperature control system, comparison of these temperatures with upper and low limiting magnitudes and delivery of heat to radiation surface when temperatures are below low magnitudes. Flight intervals at power requirement exceeding power generated by primary onboard power sources are determined. Amount of electric power consumed for temperature control of radiation surfaces is determined at the same intervals. Flight intervals for maximum possible accumulation of thermal energy on radiation surface in said zones within permissible temperatures are also determined. Expenses for radiation surface temperature control is taken into account. Before beginning of flight intervals at consumed electric power exceeding electric power generated by onboard power sources, heat is delivered to radiation surface zones which require consumption of power for their temperature control at these intervals. Delivery of heat is performed with upper limiting magnitudes of temperatures taken into account.

EFFECT: reduced loading of spacecraft power supply system due to reduced power requirement for radiation surface temperature control at retained preset temperature ranges on these surfaces.

3 dwg

Descent spacecraft // 2244665

FIELD: rocketry and space engineering; small descent spacecraft injected into orbit by ballistic missiles removed from combat duty.

SUBSTANCE: proposed descent spacecraft has case, control members for its orientation in space and units for coupling with launch vehicle. Plane of connection of descent spacecraft with launch vehicle coincides with transversal plane of symmetry of case of spacecraft and its diameter coincides in this plane with transversal diameter of launch vehicle. Descent spacecraft case is made in form of body obtained by rotation of oval around its larger axis and said plane of connection with launch vehicle coincides with plane of smaller axis of oval.

EFFECT: increased volume of descent spacecraft at limited transversal diameter of launch vehicle.

1 dwg, 1 tbl

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