RU2265559C1 - Method of launching multi-stage space launch vehicle by means of carrier-aircraft and multi-stage space launch vehicle for realization of this method - Google Patents

Abstract

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 1^st stage engine plant.

EFFECT: extended field of application.

7 cl, 5 dwg

Description

This invention relates to space technology and can be used in developing technology for launching a multi-stage space launch vehicle and in developing the design of a multi-stage space launch vehicle.

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

Demand for small spacecraft (MCAs) implies volumes of orders for their launches and expansion of this sector of the world market.

Market needs are met by launching the spacecraft from a stationary spaceport with a medium or heavy class carrier rocket as an additional payload, and a prerequisite is the availability of an order to launch a large-sized main payload.

The disadvantage of this method of launching the MCA is the limitation on their number of launches, due to the availability of orders for the removal of large-sized main payloads.

In practice, the problem of ensuring the launch of the spacecraft not only from stationary spaceports, but also from the territory of the country of the customer or owner of the spacecraft is currently being solved.

An analysis of the infrastructure of space-rocket complexes (RSC), both stationary and mobile, shows that in order to ensure the payload is brought into near-Earth space, the main parameters of the complex determine: the method of putting payload into near-Earth space, the launch method and the design of a space multi-stage launch vehicle . With all this, the RKK includes: radio engineering means of receiving external information and forming a flight program of the launch vehicle, ground handling facilities; basing system (cosmodrome, aerodrome, launch pad, prepared in engineering terms, submarine cruiser); means of transporting the launch vehicle; measuring points for receiving telemetric information about the parameters of the trajectory of the carrier rocket, separation of steps and separation of the ICA.

The task of bringing the spacecraft into the near-Earth space using the spacecraft is considered complete if the spacecraft separated from the launch vehicle at a given orbit height with the accuracy specified by the terms of the contract. Further maintenance of the ICA is provided by services not included in the infrastructure of the RCM.

An analysis of the ballistic scheme for the deducing of the spacecraft into the near-Earth space using a launch vehicle with solid-propellant propulsion systems shows that the maximum payload mass at a given orbit height can be derived using a ballistic “pause”. In this case, the separation of the MCA from the launch vehicle should optimally take place at a sufficiently large distance approximately halfway around the orbit of the MCA. This circumstance makes it difficult to bring the MCA into the near-Earth space from the customer’s territory, since it is required to place RKK elements (measuring points) outside the customer’s territory.

It is possible to solve the problem of bringing the MCA to near-Earth space from the customer’s territory using an RCM that is autonomous (independent of third parties) and mobile, which requires solving a number of interrelated problems, which include:

I. Development of a method for introducing spacecraft into the near-Earth space and development of an RSC.

II. Development of a method for launching a launch vehicle into a basing system and development of a launch vehicle.

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

Consideration of the first problem showed that for the goal, based on the choice of the appropriate parameters of the launch vehicle and the parameters of the launch vehicle, it is necessary to optimize the composition of the rocket, to solve the problem of creating an autonomous aviation space-rocket complex (ARKK) and launch the ICA from the customer’s territory . The application titled "A method for putting payload into near-Earth space using an aircraft rocket and space complex and an aircraft rocket and space complex" is dedicated to the solution of this problem for the grant of a patent for an invention, which is filed jointly and simultaneously with this application.

The consideration of the second problem showed that for this purpose it is necessary to use the advantages of a carrier aircraft with the technical characteristics of a fourth-generation fighter-interceptor of the MiG-31 type when launching a launch vehicle. The present application for the grant of a patent for an invention is devoted to the solution of this problem.

The solution to the third problem showed that for this purpose it is necessary to use a multi-stage solid-fuel rocket, because in this case, there is no need to refuel it with liquid fuel components, and this simplifies ground equipment (there is no complex refueling complex), and the rocket can be prefabricated and, as a result, high reliability, safety, convenience and ease of operation.

The third task is addressed by an application entitled "Multistage space launch vehicle" for the grant of a patent for an invention, which is submitted jointly and simultaneously with the present application.

The most important stage in the functioning of the RKK is the launch of a missile with HF. At the same time, the SCC's energy capabilities for launching spacecraft into near-earth orbits largely depend on how fully it will be possible to use the initial speed and altitude provided by the carrier aircraft at the time of the separation of the rocket from it.

As an analogue, a scheme is considered with loading the KRN into the cargo compartment of the aircraft, lifting it to a predetermined height by the aircraft, ejecting it through the tailgate of the SCC, launching the SCC by parachute and then firing it before launching the 1st stage propulsion system.

The disadvantage of this method of starting SCC is the complete loss of kinetic energy reported by SCC, and at the moment the 1st stage propulsion system is turned on, the SCC acquires a negative vertical speed, which compensates for a part of the SCC energy after starting the 1st stage propulsion system. Thus, this scheme uses, and not completely, only the increase in altitude reported by the missile to the aircraft.

Closest to the proposed method of putting the MCA into low Earth orbit is the method with the SCC planning section prior to launching the SCC 1st stage propulsion system, involving the use of a wing to maintain the necessary direction of the prelaunch missile trajectory. The application of this scheme is implemented in a space-rocket complex with an aircraft launch

- in the American Pegasus project using a non-maneuverable subsonic carrier aircraft.

The disadvantage of this method of launching the SCC is a significant weighting of the design of the first stage as a result of the installation of the wing and the difficulties associated with the layout of such a wing.

The objective of the first of the group of inventions is to expand the scope of SCC, the method of its launch while ensuring stabilization in the flight section of the safe distance from the carrier aircraft until the launch of the 1st stage propulsion system.

The problem is solved in that in the method of launching a multi-stage space launch vehicle using a launch vehicle, which consists of transporting the SCC to the launch position of the rocket, preparing the SCC for launch, lifting the SCC to a predetermined height by the carrier plane, separating the SCC from the carrier aircraft, stabilization of the SCC and the start of the propulsion system of the first booster stage, the SCC is equipped with a detachable tail fairing and folded lattice stabilizers fixed to it and transported to the collection f to the starting position with the help of a transport and operational container, overload the crane the transport and operational container from the KRC to the transport and assembly trolley, remove the removable compartments of the transport and operational container and transport the KRC to the carrier aircraft, fasten the KRC to the carrier aircraft on the airplane locks carrier, and after lifting the SCC by the carrier aircraft to a predetermined height, at the command to open the locks of the carrier aircraft, the lattice stabilizers of the tail fairing are simultaneously opened and after the calculated pause after separation of the SCC from the carrier aircraft, the tail fairing with trellised stabilizers is separated from the SCC.

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

The KRS lift is carried out on a fourth-generation MiG-31 supersonic high-altitude fighter-interceptor.

Separate the SCC from the carrier aircraft when the aircraft is cabling at an altitude of 15 ... 19 km, while the aircraft is held at a speed of 550 ... 630 m / s and a pitch angle of 15 ... 35 °.

The start of the propulsion system of the first stage is carried out after a calculated pause of 5 ... 6 s after separation of the SCC from the carrier aircraft.

The technical task of the second of the group of inventions regarding the design of the SCC is to expand the scope of its application by reducing the launch mass and ensuring stabilization in the flight section of the safe distance from the carrier aircraft until the launch of the 1st stage propulsion system.

The stated technical problem is solved in that a multi-stage space launch vehicle containing booster stages connected in series by connecting compartments and equipped with solid fuel propulsion systems, as well as a stabilization device and attachment points to the carrier aircraft, is characterized in that it is equipped with a detachable tail fairing and lattice stabilizers fixed on it, made in the form of a cylindrical panel, while the tail fairing is made with a truncated conic with a shell supported by end frames, the conical shell is closed by a spherical segment from the side of the smaller base, decaying attachments of the tail fairing to the free end of the first booster stage are located on the larger end frame, and fixation, rotation, and fastening to the trellised stabilizers are installed on the smaller end frame the lateral surface of the conical shell there are cavities that coincide in shape and dimensions with lattice stabilizers, in folded and fixed co oyanii each grid stabilizer disposed in its cavity and recessed.

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

The half-angle of the 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 m ² .

The inventive method of launching a multi-stage SCV and the claimed multi-stage SCC are united by a common inventive concept, so in the inventive method, the multi-stage SCC is launched, the starting mass of the SCC is reduced and stabilization of the SCC is ensured in the flight section of the safe distance from the carrier aircraft until the first engine is launched steps.

A specific example of the implementation of the launch method and the design of the SCC is explained in Fig.1-5.

Figure 1 shows the transport and operational container for the launch vehicle.

Figure 2 shows the transport and operational container assembly with a launch vehicle without removable compartments of the transport and operational container.

Figure 3 shows the aircraft carrier assembly with SCC.

Figure 4 shows a multi-stage SCC with a resettable tail fairing and mounted on it trellised stabilizers in the folded position.

Figure 5 shows a multi-stage SCC with a resettable tail fairing and mounted on it lattice stabilizers in the open position.

Using ground-based technological equipment, a carrier rocket equipped with fuel and pyro means, in a transport and operational container, Fig. 1, is transported from the manufacturer to the base airfield.

Transport and operational container, figure 1, is made in the form of a cylindrical pencil case and is divided by two transverse planes of the connector 1 into the front 2, middle 3 and rear 4 compartments. The front and rear compartments are made in the form of a cylindrical cup with a flat bottom and are separated by a plane 5 of the longitudinal connector with attachment points (attachment points not shown in the figures) into the upper and lower halves, the middle compartment is also divided by a plane 5 of the longitudinal connector with attachment points (attachment points on figures not shown) on the base 6 and the upper panel 7.

The mechanical connection of the rocket with the carrier aircraft is carried out by means of two fastening belts, each of which is made in the form of a bandage consisting of two half rings (not shown in the figures). Pyro-pushers are placed at the joints of the half rings, with the help of which half of the belts are pushed apart when they are reset after separation of the SCC from the aircraft. The front mounting belt is equipped with removable power frames. Power frames are connected to the grips on the plane. During ground operation, the power frames are removed (not shown in the figures) to place the rocket in the transport and operational container.

A mechanical coupling kit (not shown in the figures) ensures that the KRC is placed on the lodges of the transport and operational container, the KRC is secured in the container from axial movement and twisting during transportation, workshop loading and unloading operations are carried out with the KRC, the KRC is fixed to the aircraft suspension and the mechanical elements are separated communications from SCC after undocking from the aircraft.

At the base airfield, a transport and operational container with a launch vehicle is installed and fixed on a transport and assembly trolley (not shown in the figures). On the transport and assembly trolley, the KVN is transported within the technical zone of the airfield and the work preceding the KVN suspension under the carrier aircraft is carried out. Transportation of the transport and assembly trolley with SCC for the plane is carried out after dismantling the removable elements of the transport and operational container: the front and rear compartments, the upper panel of the middle compartment, while the SCC is fixed to the base 6, figure 2, the middle compartment.

When KRS is hung to an airplane, the transport and assembly trolley is hung on jacks (not shown in the figures). After the KRC is suspended from the aircraft, the reliability of the lock closure is checked, the KRC is unfastened from the locks of the transport and assembly trolley and the trolley frame is lowered to ensure that the KRC is hung on the airplane locks (not shown in the figures).

Aircraft carrier 8 with SCC 9 is shown in Fig.3.

To reduce the aerodynamic drag of the SCC on the suspension under the aircraft, the SCC contains a tail fairing 10, Fig. 4, with folded lattice stabilizers 11.

To stabilize the movement of SCC after separation from the carrier aircraft, the lattice stabilizers 11, Fig. 5, are forced to open.

After the SCV has been lifted by the carrier aircraft to a predetermined height by the command to open the locks of the carrier aircraft (not shown in the figures), the lattice stabilizers of the tail fairing are simultaneously opened and, after a calculated pause, before starting the engine installation of the 1st accelerating stage, the tail fairing with lattice stabilizers is separated from SCC.

Schematic design decisions on the system for separating SCC from the carrier aircraft and the launch pad of SCC are selected based on the need to turn on the 1st stage SCC propulsion system at a distance of more than 60 m from the plane to ensure its safety (~ 5 s after separation of SCC). The KVN static stability in an uncontrolled flight section after separation from the aircraft until the 1st stage propulsion system is turned on is ensured by the disclosed lattice stabilizers located on the KVN tail fairing discharged before turning on its engine.

Conditions for the separation of SCC from CH:

- height - 15 ... 19 km;

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

- transverse overload of the aircraft at the time of separation - +2.4 ... + 2.5.

Significant transverse overload of the aircraft allows without the use of special devices to provide shock-free separation of SCV. The command to open the lattice stabilizers is given upon the rupture of the mechanical connection between the SCC and the aircraft.

Under these conditions, 0.3 s after separation of the SCC from the carrier aircraft, the distance between them will exceed 1 m, which allows by this time to complete the disclosure of trellised stabilizers. After the stabilizers are opened, the KRC has a positive margin of static stability and performs weakly damped angular oscillations relative to the center of mass. With a minimum allowable margin of static stability of 7-8%, the effective area of lattice stabilizers will be 2.8 - 3.0 m ² . After separation from the aircraft, the on-board KRS control system according to the current parameters of movement, based on the conditions of departure from the aircraft to the required distance and the possibility of subsequent flight control, determines the moment of dumping the tail fairing and the start of the propulsion system of the 1st stage of KPH.

After separation of the tail fairing from the SCC, a start is made.

A specific embodiment of a multi-stage space launch vehicle is explained in FIGS. 1-5.

To implement the method of launching multi-stage SCC using a carrier aircraft, a space launch vehicle has been developed, which is shown in FIG. 4.

Based on the high value of the required final speed (-8000 m / s) required for the SCV, taking into account the fact that part of this speed is reported before the rocket starts by plane, a scheme with three marching stages and apogee stage 12 is selected for the SCC, Fig. 4, with a solid fuel propulsion system 13.

To reduce the aerodynamic drag of the rocket on the suspension under the plane and to give the rocket static stability in the area of uncontrolled movement from the moment it is separated from the plane to start the engine when removed at a safe distance from the plane, the rocket contains a tail fairing 10, figure 4, with folded lattice stabilizers 11 forcibly disclosed after the discharge of SCC from the aircraft. The tail fairing is separated immediately before the start of the propulsion system of the 1st stage of the SRV (not shown in the figures).

The tail fairing is made in the form of a truncated conical shell, supported by end frames. From the side of the smaller base, the conical shell is closed by a spherical segment 14, Fig. 4, decaying nodes (not shown) of the fastening of the tail fairing 10 to the free end of the first booster stage are placed on the larger end frame. On the smaller end frame, fixation, rotation and fastening nodes (not shown in the figures) of lattice stabilizers 11 are installed. On the side surface of the conical shell there are cavities 15, Fig. 5, matching the shape and dimensions of the lattice stabilizers 11. In the folded and fixed state, each the trellis stabilizer is placed in its cavity and recessed.

Thus, it has been shown that the proposed method for launching a multi-stage SCC using a carrier aircraft and the developed design of a multi-stage SCC can reduce the launch mass of the SCC and take advantage of the carrier aircraft with the technical characteristics of a fourth-generation fighter-interceptor of the MiG-31 type when launching a launch vehicle.

Claims (7)

1. The method of launching a multi-stage space launch vehicle (SCV) using a launch vehicle, which consists in transporting the SCC to the launch position of the rocket, preparing the SCC for launch, lifting the SCC to a predetermined height by the carrier plane, separating the SCC from the carrier aircraft, stabilizing the SCC and the start of the propulsion system of the first booster stage, characterized in that the SCC is equipped with a detachable tail fairing and folded lattice stabilizers fixed to it and transported assembly to the starting position with transport and operational container, crane overload the transport and operational container from the SRV to the transport and assembly trolley, remove the removable compartments of the transport and operational container and transport the SRV to the carrier aircraft, fasten the SRC to the carrier aircraft on the locks of the carrier aircraft, and after lifting the SCC by the carrier aircraft to a predetermined height by the command to open the locks of the carrier aircraft simultaneously open the lattice stabilizers of the tail fairing and after the calculated pause d by separating the SCC from the carrier aircraft, the tail fairing with trellised stabilizers is separated from the SCC. 2. The method according to claim 1, characterized in that the start of the first-stage propulsion system is carried out after a calculated pause of 5-6 s after separation of the SCC from the carrier aircraft. 3. The method according to claim 1, characterized in that the SCC is separated from the carrier aircraft when the aircraft is being cabrated at an altitude of 15-19 km, while the aircraft is held at a speed of 550-630 m / s and a pitch angle of 20-40 °. 4. A multistage space launch vehicle containing booster stages connected in series by connecting compartments and equipped with solid-fuel propulsion systems, as well as a stabilization device and attachment points to the carrier aircraft, characterized in that it is equipped with a detachable tail fairing and mounted on it lattice stabilizers made in the form of a cylindrical panel, while the tail fairing is made with a truncated conical shell, reinforced by end frames By the way, from the side of the smaller base the conical shell is closed by a spherical segment, on the larger end frame there are decaying attachments of the tail fairing to the free end of the first booster stage, and on the smaller end frame there are fixation, rotation and fastening units of the grating stabilizers, on the side surface of the conical shell cavities matching in shape and dimensions with trellised stabilizers, in the folded and fixed state, each trellised stabilizer is accommodated in its cavity and drowned. 5. The launch vehicle according to claim 4, characterized in that the half-angle of the conical shell of the tail fairing is 15-20. 6. The launch vehicle according to claim 4, characterized in that the radius of the spherical segment of the tail fairing is 200-300 mm 7. The launch vehicle according to claim 4, characterized in that the effective area of trellised stabilizers is 2.5-3.5 m

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