Method of injection of payload into near-earth space by means of aircraft rocket space complex and aircraft rocket space complex for realization of this method

FIELD: space engineering.

SUBSTANCE: proposed method includes joint assembly of payload and launch vehicle for forming space launch vehicle which is equipped with apogee stage with solid-propellant engine plant. Carrier-aircraft is coupled with space launch vehicle and launch vehicle is raised by this aircraft to preset altitude, then launch vehicle is separated and solid-propellant engine plants of three boost stages are started in succession; launch vehicle is injected into preset near-earth orbit and payload is separated from launch vehicle at preset point of trajectory in preset direction. In the course of flight of launch vehicle upon discontinuation of operation of engine plants of boost stages and completion of first boost leg, ballistic pause is performed at motion of space launch vehicle over ballistic trajectory at climbing the required altitude of orbit. Upon completion of ballistic pause at second boost leg engine of apogee stage is started and space launch vehicle is injected into preset near-earth orbit at respective velocity increment and compensation of error during operation of boost stages. Aircraft rocket space complex includes 1st class aerodrome, carrier-aircraft and space launch vehicle. Masses of boost and apogee stages are selected at definite ratio. Provision is made for transportation container for delivery of space launch vehicle to aerodrome. Telemetric information measuring and tracking points are located on aeroplanes; they are made in form of mobile radio unit for reception of external information.

EFFECT: reduction of distance from launch site of space launch vehicle to point of separation of payload.

18 cl, 11 dwg

 

This invention relates to space technology and can be used in the development of technology output payload into near-earth space and the development of aviation and rocket-space complex (RSC).

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 owner of the MCA.

Infrastructure analysis RCM, both stationary and mobile, shows that to ensure the withdrawal of polezno the cargo in 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 consists of Radiotechnical means of receiving external information and formation flight, launch vehicle, ground support, system-based spaceport, airfield, launch pad, trained in engineering terms, submarine), means of transportation booster, as well as the measuring points of reception of telemetry information about the parameters of the trajectory of the rocket branch of the levels and branches of the MCA.

The objective of the RCM at the conclusion of the MCA on the near-earth space is considered to be fulfilled 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". When this branch of the MCA from the booster ideally should take place at a great distance, approximately provide motion along the orbit of the ICA. This circumstance complicates the OS which may serve as a basis conclusion of satellites in near-earth space with the customer by means of RCMS, since you want to put the 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.

The solution of the first question of the complex issues devoted to the present invention.

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 with the launch of a space rocket (SCC). The solution to this problem is devoted to an application entitled "Method of starting a multi-stage space launch vehicle using aircraft-carrier and 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 solution of the third problem was shown that for postavlennaya you must use a 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 closest to the proposed method output ľa in low earth orbit is the way of the conclusion in the near-earth space ICA using an RCM-based transportation to the launch pad MCA and booster in containers, their joint coupling, the start of SCC, collecting telemetry data measuring points tracking and processing information about the trajectory parameters of SCC, the Department of stages and the separation of the payload and the Department of MCA from SCC in a given direction ľa (EN 93048257 C1, 1997.01.27).

Closest to the proposed RCM is complex, containing the booster with booster stages, with the propulsion system on solid fuel, MCA, ground maintenance, electronic complex, measuring receiving and processing the TCI telemetry data about the trajectory of the rocket, offices stages and ICA (see ibid.).

The disadvantages of this method o payload into near-earth space and ALAC are in need of large areas to accommodate its infrastructure.

The objective of the first group of inventions is to optimize the composition of the RCM-based selection of relevant motion parameters of SCC.

The problem is solved in that in the output payload into near-earth space using airborne missile and space complex, according to which is transported to the airfield the aircraft carrier, booster, payload, maintenance and preparation of the carrier aircraft and rocket to fly, radio-technical means for the formation flight, the carrier aircraft and launch vehicle, and ground equipment, carry out joint Assembly payload and booster with the formation of a space rocket, connect the aircraft carrier with SCC, raise KRN aircraft carrier to a predetermined height, separated KRN from the carrier aircraft, consistently launch propulsion systems solid fuel three booster stages, the output of SCC on a given orbit, and at the clearing point of the trajectory separates the payload from SCC in a given direction, is that pre-determine the coordinates of the location of the airfield and the parameters of the trajectory KRN, establish area offices and downs spent all its stages and branches of the payload and telemetry data gathering about the trajectory parameters of SCC, the separation of the individual stages and the separation of the payload to produce the measuring points tracking and processing information, exercise equipment KRN apogee stage with the propulsion system on solid fuel, is delivered to the airfield booster loaded with fuel and procreative, operating freight container, during the flight of SCC after propulsion booster stages and completion of the first upper area carry out ballistic break with the movement of SCC on a ballistic trajectory and set the desired height of the orbit, completion ballistic pause on the second boost phase, start the motor installation apogee stage, and on a given orbit deduce SCC with a corresponding increment of speed and compensate for the errors booster stages, with characteristic speed apogee stage set in the range of 14...16% of the total rate for all stages of SCC, and telemetry data gathering exercise measuring points placed on the aircraft, and mobile radio authority receiving the external and the formation.

The task promote private significant features of the invention.

Provide the inferred mass of the spacecraft 50-150 kg, orbital inclination 45,8-90°, the height of the orbit 280-1000 km, and the errors reach orbital altitude and orbital inclination are respectively 1.8% and 0.7 percent.

The duration of the ballistic pause set in the range of 300...800 C.

The separation of the payload from apogee stage is conducted at a distance of 3300...3700 km from the launch site.

KRN perform with a launch mass of 10...11 t, length 10...11 m and a maximum diameter of 1.3...1.4 m, and the center of gravity of SCC have a distance of 6...6.3 m from the toe of the head fairing.

For the exhaust of the propulsion system of the third booster point fall along a ballistic trajectory set at a distance of 5000...10000 km from the launch site.

During the transportation of the booster to the airfield in operating freight container support a temperature range from +5°C to +30°C.

Transportation of the booster vehicle maintenance container to the airfield is carried out in isothermal car or on the plane An-124.

The task of the second group of inventions is the creation of Autonomous aircraft missile and space complex, allowing besitkas the th running ľa fully solid-fuel booster weighing up to 11 kg per airfield 1st class.

The technical problem is solved by the fact that in aviation and rocket-space complex containing the airfield 1st class aircraft carrier, the booster and payload, forming a collection space carrier rocket with three booster stages, equipped with propulsion units for solid fuel, maintenance and preparation of the carrier aircraft and rocket to fly, radio-technical means for the formation flight, the carrier aircraft and launch vehicle, ground equipment, and instrumentation items tracking and processing of telemetry data about the trajectory parameters of SCC, the Department of individual steps and the separation of the payload, KRN equipped apogee stage with the propulsion system on solid fuel, and the mass of the propulsion units mDNI PC, mDN II PC, mDN III PCand mDo ACaccordingly, the first, second, third and apogee stages selected ratios:

mDo I PC/mDN II PC=1,7 1,9...,

mDN II PC/mDN III PC=2,1 2,4...,

mDN III PC/mDo AC=4,0...4,2,

for delivery to the airfield booster loaded with fuel and procreative made operating freight container, and measuring items tracking and processing of telemetry data, saving the NY jets and made in the form of a mobile radio authority receiving external information.

The task promote private significant features of the invention.

Operating freight container is made in the form of pencil case has a cylindrical shape and is divided in two transverse planes of the connector with attachment points on front, middle and rear compartments, the front and rear compartments made in the form of a cylindrical Cup with a flat bottom and separated by a longitudinal plane of the connector with attachment points on the upper and lower halves, the middle compartment is also divided by a longitudinal plane of the connector with attachment points on the base and the top panel, the base of the middle compartment is made in the form of a cylindrical panel with two cradles, on the outer side of the cradle is made mountings of the container to vehicles and components for crane overload container.

SCC is equipped with a set of mechanical connection of SCC with suspension of the carrier aircraft, containing two bandages, consisting of upper and lower half-rings, joints semicircles connected by hinge mechanisms with spreading attachment points, and each upper semi-fixed force frame, with the power frame is made removable.

The measuring points of the tracking and processing of information placed on two planes the IPA IL-76MD.

As a means of transportation means of the ground equipment used by the aircraft IL-76MD.

As the carrier aircraft used supersonic high-altitude fighter-interceptor fourth generation MiG-31.

The inventive method and the inventive aircraft RCM United General inventive concept, the inventive method discharge payload into earth orbit using the infrastructure of Autonomous mobile ALAC.

A specific example of the method and composition of the aircraft RCM illustrated in figure 1-11.

Figure 1 shows the composition of the aviation ALAC.

Figure 2 shows the aircraft-carrier Assembly with SCC.

Figure 3 presents the parameters of the trajectory of SCC and a circuit receiving the telemetry data of the measuring points placed on the aircraft.

Figure 4 shows the space multistage SCC with motors on solid fuel.

Figure 5 shows a transport and operational container for the rocket.

Figure 6 shows operating freight container Assembly with LV.

Figure 7 shows the section AA figure 6.

On Fig depicted SCC in operating freight container after the removal of its removable parts.

Figure 9 depicts a front waistband with detachable power RA is Oh.

Figure 10 shows the back belt with detachable power frame.

Figure 11 shows the hinge mechanism from the disintegrating fixation half-rings of the binder.

Aviation ALAC, figure 1, which is used to launch satellites in near-earth space, consists of a booster I, aircraft carrier II, mobile radio engineering complex reception of external information and formation flight program aircraft carrier (SN) and the rocket III, the mobile complex ground support equipment IV, basic airfield 1st class V, means of transportation of the mobile complex of ground technological equipment VI, airborne command and measurement points VII.

Using ground-based technological equipment is a complex of measures on the prelaunch KRN equipped apogee stage with the propulsion system on solid fuel, starting from sending KRN filled with fuel and procreative in operating freight container manufacturers, before the release of SN with SCC on the runway for the next takeoff. Aircraft-carrier 1 with comic booster 2 is shown in figure 2.

After joining the MCA to the booster and LV electrical tests of the control system of SCC and enter the flight task raise KRN plane-wear is eating at a predetermined height, separate KRN from the carrier aircraft. When removing the SCC at a safe distance from the aircraft carrier, start the motor the installation of solid fuel 1st booster, which corresponds to point a, figure 3, on the trajectory of the SCC. Department 1st booster 3 and the start of the propulsion system solid fuel 2nd booster corresponds to the point B, figure 3, on the trajectory of the SCC. Department 2nd booster 4, the Department head fairing 5 and launch propulsion systems solid fuel 3rd booster corresponds to the point C, figure 3, on the trajectory of the SCC. Department of the 3rd booster 6 and the beginning of the movement apogee stage 7, figure 3, with a long ballistic pause corresponds to the point D, figure 3, on the trajectory of the SCC. Point D in figure 3 corresponds to the end of the first upper portion of the trajectory of movement of SCC.

Upon completion of the ballistic pause on the second boost phase, start the motor installation apogee stage, point E in figure 3, with the characteristic speed of apogee stage set in the range of 14...16% of the total rate for all stages of SCC.

Spend the velocity increment apogee stage and compensation of errors in the work of booster stages. At the clearing point of the trajectory of the SCC, point E in figure 3, separate ľa from the PH in a given direction.

In connection with the characteristic of SCC with motors on solid fuel is relatively small (less than minutes) work times propulsion units, each of the upper (main) levels, as well as due to the lack of practical possibility of multiple switching and level control thrust solid-propellant rocket engines height end area of propulsion stages in their continuous sequential involvement is no more than 150-250 km

In the absence of additional measures this value limits the maximum height of a circular or elliptical orbit perigee output ľa insufficient for prolonged existence of the MCA due to the heavy braking its movement by the Earth's atmosphere.

To enable removal of the ICA on the orbit with a perigee of 300 to 1000 km for SCC provides functional layout with long ballistic "pause"in which is set the desired height of the orbit. Pause starts at the end of the first upper area - point D, figure 3, and ends before the beginning of the second accelerating section to point E, figure 3, getting started apogee propulsion stage.

Part of apogee stage along with the propulsion system includes a development engine of small thrust, ensuring the th accuracy needed to launch the ICA into orbit, and vasoreactivity orientation system designed to hold the angular position of SCC in the area of "pause" in the limits required by the conditions of health of the onboard equipment.

Accepted for SCC in the present invention is characteristic speed apogee stages 14-16% of the total rate for all stages of SCC exceed the same amount for theoretically optimal scheme with darshanam upon completion Polovinka transition elliptical orbit and provides:

the approximation points branch of the MCA and booster stages to the launch area 3-5 times, which ensures reliable reception of information transmitted from the Board of the SCC measuring points;

- reduction of three times the duration of the ballistic "pause" from -45 to -15+20 min, thereby facilitating mass sources of power for onboard systems, GRSO by an amount equal to 10-15% of the mass output of the AC, as well as improving the accuracy of the SPACECRAFT;

- removing spent the last sustainer stage on a ballistic trajectory with the removal of the point of falling to - 10000 km, and not on an elliptical orbit with an unpredictable impact area.

The calculation of the distances separating stages of SCC and MCA showed:

1. Department 1st booster is at a distance of S=50 km from the starting point at t=45 s from the starting time on high is e N=36 km, the point a on the trajectory of the SCC, 3.

2. Department 2nd booster is at a distance S=205 km from the launch site, at t=100 C from the starting time, at a height of H=88 km point on the trajectory of the SCC, 3.

3. Department of the 3rd booster and the beginning of a long ballistic pause occurs at the distance S=452 km from the launch site, at t=151 from the starting time, at a height of H=141 km, point D on the trajectory of the SCC, 3.

4. Long ballistic pause ends at a distance S=3111 km from the launch site, at t=600 seconds from the start time, at a height of H=426 km, point E on the trajectory of the SCC, 3.

5. The Department of MCA from SCC occurs at the distance S=3441 km from the launch site, at t=648 with from the start time, at a height of H=426 km G point on the trajectory of the SCC, 3.

This example shows that using the proposed method output payload into near-earth space Department of MCA from SCC can be performed at a distance of 3700 km from the starting point, and the branch of the 1st, 2nd and 3rd booster stages, respectively, at distances of 50 km 205 km, 452 km from the start of the SCC. This in turn provides the possibility of receiving information from SCC on its trajectory, and about the separation of all of its steps and ICA using the measurement points placed on the planes indicated by the letters X and Y in figure 3, within which territorii customer. In this takeoff of aircraft with measuring points is made with the base airfield customer. After establishing that the Department of MCA from SCC by means of Autonomous RCM task ALAC output of satellites in near-earth space is performed.

Thus, the entire infrastructure of the RCM, including the measuring points may be placed at the customer's site in the Autonomous aircraft ALAC using the developed method o MCA in near-earth space.

A specific example of aviation and rocket-space complex is shown in Figure 1-7.

Aviation ALAC, figure 1, which is used to launch satellites in near-earth space, consists of a booster I, aircraft carrier II, mobile radio engineering complex reception of external information and formation flight program aircraft carrier (SN) and the rocket III, the mobile complex ground support equipment IV, basic airfield 1st class V, means of transportation of the mobile complex of ground technological equipment VI, airborne command and measuring points VII.

To implement the method o payload into near-earth space using Autonomous aircraft RCM developed SCC, is shown in figure 4.

On the basis of a high value pot is ebnoy final speed (-8000 m/s), necessary for SCC, taking into account the fact that part of this speed is reported before the start of SCC aircraft selected for SCC scheme with three lively marching (upper) levels, which contains apogee stage 8, figure 4, with the propulsion system on solid fuel 9.

When a three-stage version of the rocket due to the mass of the propulsion system 10, 4, 1 booster gives the level of the velocity head at the division 1-St and 2-nd stages 1300 kg/m2allowing the use of 2-Oh booster 11, figure 4, a slight rotary nozzle 12 with the angle of deviation 3°, 4.

Based on the results of optimization calculations of the distribution of fuel in the propulsion units propulsion stages of SCC, the mass of the propulsion units (DU) mDNI PC, mDN II PC, mDN III PCand mDo AC1st, 2nd, 3rd and apogee stages, respectively, 10, 13, 14, and 9, figure 4, selected ratios:

mDo I PC/mDN II PC=1,7 1,9...,

mDN II PC/mDN III PC=2,1 2,4...,

mDN III PC/mDo AC=4,0...4,2.

This allows the use of propulsion systems 10, 13, 14, and 9, figure 4, to realize the value of the characteristic speed apogee stages 14-16% of the total rate for all stages of SCC, the Department of MCA from SCC conducted at a distance of 3700 km from the starting point. Department of the 1st, 2nd and 3rd accelerating the stupa is it carried out respectively at distances of 50 km, 205 km, 452 km from the start of the SCC. Receiving information from the SCC on its trajectory, and about the separation of all of its steps and the MCA carried out with the use of measurement points placed on the planes indicated by the letters X and Y in figure 3, within the territory of the customer. Takeoff aircraft with measuring points make with the base airfield customer, but the fact of the Department of ICA from SCC carried out by means of Autonomous ALAC. The objective of the RCM at the conclusion of satellites in near-earth space is performed.

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 removed to a safe distance from the aircraft, the missile contains a tail fairing 15, figure 4, with folded lattice aerodynamic stabilizers 16, forcibly disclosed after reset KRN from the plane (figures not shown). Tail fairing is separated immediately before the start of the propulsion system of the 1st stage of SCC (figures not shown).

For the entire period of the operating cycle, beginning with a boot to the factory and ending with a snap of SCC to the plane of the carrier, developed transport and operational container 5.

Transport-ek is operating the container, 5, in the form of pencil case has a cylindrical shape and is divided in two transverse planes of the connector 17 and 18 with fasteners (fasteners on the figures not shown) on the front 19 average 20 and 21 rear compartments. Front and rear compartments made in the form of a cylindrical Cup with a flat bottom and separated by a longitudinal plane 22 of the connector with fasteners (fasteners on the figures not shown) on the upper and lower halves, the middle compartment is also divided by a plane 22 of the longitudinal connector with fasteners (fasteners on the figures not shown) on the base 23 and the top panel 24. The base 23 of the middle compartment is made in the form of a cylindrical panel with the cradle 24 and 25. On the outer side of the cradle is made sites for crane overload container 26, 5 and 7, the attachment points of the container to the transport means 27, 7.

Due to the presence of a layer of insulation within the transport and maintenance of the container and sealing his joints (figures not shown) when connecting refrigeration and heating unit provided for SCC temperature and humidity conditions (from +5 to +30°C, from +10 to +25°in the area of the instrument compartment). When the unit is insulated container ensures the preservation of the specified temperature of the product within 4 hours at the temperature of environment is about air -40° C and the initial temperature of the product +25°and for 30 hours at an ambient temperature of +40°and the initial temperature of the product +10°C.

Cradles 24 and 25, 5 and 6, the base 23 of the middle compartment 20 are used for laying of SCC and published in the areas of power frames SCC 28 and 29, 6. On the outer side of the cradle is made of the support surface 30, Fig.7, for installation and fastening of the container with SCC on vehicles (vehicles figures not shown). Top panel front cover, middle and posterior compartments are removable. On the top panel of the middle compartment (figures not shown) made a hatch for access to the electrical connectors of the docking SCC.

In the operation of transport and operational container with SCC as possible elementwise, and simultaneous removal of the front compartments 19 and 21 rear and upper halves of these compartments together with the upper panel of the middle compartment 24.

Consolidation of SCC in operating freight container in the axial and radial directions during transportation, as well as from twisting by means of a contact zone (from kit mechanical connection SCC) with the groove on the front lodgement (figures not shown), and install the tie ribbon 31, 7, attached to each lodgement. Tape removed at the stage of preparation the Cai to the connection of SCC with the aircraft carrier.

SCC enshrined in operating freight container, with removed front and rear compartments in Assembly with the base 23 and the clamping strips 31 shown in Fig.

SCC is equipped with a set of mechanical connection of SCC consisting of two belts fastening: front, Fig.9, and back, figure 10.

The mechanical connection kit 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 before start after separation from the aircraft.

On SCC provides grooves for the installation of zone (figures not shown).

Each belt is made in the form of a bandage comprising upper and lower semicircles: front belt 32 and 33, respectively, figure 9, and the rear belt 34 and 35, respectively, figure 10. The joints of the half rings are connected by two hinge mechanisms 36, figure 9 and figure 10. Hinge mechanism 11, is made from decaying node connection 37, 11, for example, in the form of peerbolte, with which the ring is separated from SCC after its separation from the carrier aircraft.

To the upper semicircle 32 front belt, Fig.9, fixed power frame 38 and to the upper semicircle of the back of the belt 35, IG, fixed yoke 39 for interfacing with locks pendants aircraft carrier.

The bottom half of the zones have two screw stop 40 and 41 located at an angle of 90°. The stops are designed to prevent twisting of the belts relative to the rocket due to contact with the stoppers located on the landing frame KRN (figures not shown).

The power frame of the front zone and the rear stirrup belt have attachment points (figures not shown)that allows you to remove them for ease of placement of SCC in operating freight container during ground operation.

Ground technological equipment SCC is designed to perform complex actions on the prelaunch, starting from sending SCC and MCA with manufacturers to exit SN with SCC on the runway for the next takeoff.

The composition of the ground processing equipment includes:

Traffic machine designed for temporary storage and transportation operating freight container with SCC within the airfield, which is a self-propelled trailer. Loading operating freight container with SCC of the transport vehicle is a crane of sufficient capacity certified to work with a bit of cargo. Towing transport machines producing the automobile tractor type KrAZ-255B, with appropriate couplings for connecting the brake system and electrical equipment transportation machines.

Transport and mounting trolley is designed for transportation of SCC within the technical zone of the aerodrome, fixing KRN when performing the finish of the scope of work prior to the suspension of SCC under the plane - test works, dismantling compartments operating freight container, fastening KRN when performing the finish of the scope of work prior to the suspension of SCC under the aircraft, transportation of SCC to SN and install it under the plane. Loading SCC with operating freight container on the cradle transport and Assembly of trucks by road crane with a capacity of >20 t certified to work with a bit of cargo. Transportation transport and mounting trolley with SCC under the plane (hardpoints) after dismantling the crane removable elements operating freight container front and rear compartments, the top panel of the middle compartment. Transportation is carried out using autotrace from the tail of the plane. Under suspension of SCC to the plane of the transport and mounting trolley posted on jacks. Consistent inclusion of certain systems operator using Wynona what about the remote control carries rise KRN, achieves precise alignment of the suspension units SCC and castles of the aircraft. It is then a further rise in SCC at minimum speed jacks until lockout. 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 aircraft. Followed by the lowering of the transport and Assembly of the trolley on wheels, coupling it with the truck and it rolled out from under the plane.

Railway refrigerator car (still ALIVE) provides transportation of SCC in operating freight container offline on the Railways in the composition of freight trains and individual locomotive at speeds up to 120 km/h Systems and equipment LIVES provide: placement and consolidation of operating freight container with SCC on the retractable frame of the cargo compartment; automatic TVR inside the cargo compartment from +5 to +35°C and relative humidity not more than 80% at ambient temperatures from -50 to +50°C; detection and elimination of fire in the offices of the car; remote monitoring of power supply systems, temperature control, fire protection and alarm from the car maintenance; snabzheniem car from its own power supply (diesel generator set and the car battery or from car maintenance; two-way telephone communication with car maintenance.

- To maintain temperature and humidity conditions of SCC when performing work on the airfield used aggregate drying, heating and cooling. Technical characteristics of the unit can provide not only temperature and humidity conditions of SCC in operating freight container, but also to maintain, if necessary, thermal regime in the area of jobs or equipment and technical equipment of the complex. The equipment is installed in a special body-van. The control unit automated or remote-manual.

Kit Assembly and maintenance headunit (adapter Assembly, MCA and nose fairing - head unit) with SCC, transportation headunit in place to work with MCA and back, docking and undocking of the adapter nose fairing, mounting the unit on the SCC.

For snapping KRN to the plane of the carrier and carrying out joint inspections of systems of SCC and navigation equipment of the aircraft carrier at the aerodrome mount easily constructed hangar. At the time of execution of the work in the hangar there are: location for transport and mounting trolley with SCC in operating freight container; space to accommodate necessary the equipment kit Assembly and maintenance of SCC, body-trailer with a radio apparatus and equipment, providing the necessary training aircraft; a workstation for docking SCC and CH. The hangar has floor dimensions not less than 35×20 m and the height from the floor to the ceiling along the axis of the hangar is not less than 7.5 m hangar Floor is solid and flat floor that can withstand specific pressure from the wheels of the transport and Assembly of the truck chassis and the carrier aircraft. The hangar is equipped with: mechanical sliding or hinged gates with curtain and sizes 4,5×6 m and 16×7 m for the passage of transport and installation trucks and chassis of the aircraft carrier, respectively; temporary or permanent entries (at least two) 50 kW, 380/220 V, 50 Hz from uninterruptible power systems aerodrome; lighting system; maintain the desired temperature (temperature from +15 to +25°); ventilation; fire extinguishing means.

Measuring points tracking and processing the telemetry data is hosted on two aircraft IL-76MD and transported to the base airfield.

Ground technological equipment transport aircraft IL-76MD.

1. The output payload into near-earth space using airborne missile and space complex, according to which is transported to the airfield the aircraft carrier, Aceto-media payload, maintenance and preparation of the carrier aircraft and rocket to fly, radio-technical means for the formation flight, the carrier aircraft and launch vehicle, and ground equipment, carry out joint Assembly payload and booster with the formation of a space rocket (SCC), join the aircraft carrier with SCC, raise KRN aircraft carrier at a predetermined height, separated KRN from the carrier aircraft, consistently launch propulsion systems solid fuel three booster stages, the output of SCC on a given orbit, and the estimated point of the trajectory separates the payload from SCC in a given direction, while pre-determine the coordinates of the location of the airfield and the parameters of the trajectory of SCC, establish area offices and downs spent all its stages and branches of the payload and telemetry data gathering about the trajectory parameters of SCC, the separation of the individual stages and the separation of the payload to produce the measuring points tracking and processing information, wherein the SCC equip apogee stage with the propulsion system on solid fuel, is delivered to the airfield booster loaded with fuel and procreative, transport is about operating the container, during the flight of SCC after propulsion booster stages and completion of the first upper area carry out ballistic break with the movement of SCC on a ballistic trajectory and set the desired height of the orbit, upon completion of the ballistic pause on the second boost phase, start the motor installation apogee stage and on a given orbit deduce SCC with a corresponding increment of speed and compensate for the errors booster stages, with characteristic speed apogee stage set in the range of 14-16% of the total rate for all stages of SCC, and telemetry data gathering exercise measuring points placed on the aircraft, and mobile radio authority reception of external information.

2. The method according to claim 1, characterized in that provide the inferred mass of the spacecraft 50-150 kg, orbital inclination 45,8-90°, the height of the orbit 280-1000 km

3. The method according to claim 1, characterized in that provide error reach orbital altitude of 1.8%, the error of the inclination of the orbit of 0.7%.

4. The method according to claim 1, characterized in that during the transportation of the booster to the airfield in operating freight container support the specified temperature range.

5. The method according to claim 4, characterized in that the transport operation container to keep the temperature from 5 to 30° C.

6. The method according to claim 4, characterized in that the transportation of the booster vehicle maintenance container to the airfield is carried out in isothermal wagon.

7. The method according to claim 4, characterized in that the transportation of the booster vehicle maintenance container to the airfield carried out on the aircraft type An-124.

8. The method according to claim 1, characterized in that the duration of the ballistic pause set in the range of 300-800 C.

9. The method according to claim 1, characterized in that the separation of the payload from apogee stage is conducted at a distance 3300-3700 km from the launch site.

10. The method according to claim 1, characterized in that the SCC perform with a launch mass of 10-11 t, length 10-11 m and a maximum diameter of 1.3-1.4 m, and the center of gravity of SCC have on distance 6-6,3 m from the toe of the head fairing.

11. The method according to claim 1, characterized in that the spent motor third set booster point fall along a ballistic trajectory set at a distance of 6000-10000 miles from the launch area.

12. Aviation and rocket-space complex containing the booster and payload, forming a collection space booster (SCC) with three booster stages, equipped with propulsion units for solid fuel, airfield 1st class aircraft-carrier, means t the ical service and preparation of the carrier aircraft and rocket flight radio AIDS to the formation flight, the carrier aircraft and launch vehicle, ground equipment, and instrumentation items tracking and processing of telemetry data about the trajectory parameters of SCC, the separation of the individual stages and the separation of the payload, wherein the SCC equipped apogee stage with the propulsion system for solid fuel, with mass mDNI PC, mDNII PC, mDNIII PCand mDNACaccordingly, the first, second, third and apogee stages selected ratios:

mDo I PC/mDN II RS=1,7 1,9...,

mDN II PC/mDN III PC=2,1 2,4...,

mDNIII PC/mDNAC=4,0...4,2,

for delivery to the airfield booster loaded with fuel and procreative made operating freight container, and measuring items tracking and processing of telemetry data placed on the aircraft, and made in the form of a mobile radio authority receiving external information.

13. The complex according to item 12, characterized in that the transport services container in the form of pencil case has a cylindrical shape and is divided in two transverse planes of the connector with attachment points on front, middle and rear compartments, front and rear use is in the form of a cylindrical Cup with a flat bottom and separated by a longitudinal plane of the connector with attachment points on the upper and lower halves, the middle compartment is also divided by a longitudinal plane of the connector with attachment points on the base and the top panel, the base of the middle compartment is made in the form of a cylindrical panel, curved ends which are fixed to the cradle, on the outer side of the cradle is made mountings of the container to vehicles and components for crane overload container.

14. The complex according to item 12, characterized in that the SCC is equipped with a set of mechanical connection of SCC with suspension of the carrier aircraft, containing two bandages, each of which consists of upper and lower half-rings, joints semicircles connected by hinge mechanisms with spreading attachment points, and each upper semi-circle attached to the power frame.

15. Complex 14, characterized in that the power frame is made removable.

16. The complex according to item 12, characterized in that the measuring points of the tracking and processing of information placed on two aircraft IL-76MD.

17. The complex according to item 12, characterized in that as a means of transportation means of the ground equipment used aircraft IL-76MD.

18. The complex according to item 12, characterized in that as the carrier aircraft used supersonic high-altitude fighter-interceptor fourth generation MiG-31.



 

Same patents:

Lamellar sheathing // 2265520

FIELD: mechanical engineering; production of constructions made out of composite materials of high-precision hardware products of space and land application.

SUBSTANCE: the invention is pertaining to the field of mechanical engineering, in particular, to production of constructions made out of composite materials for the high-precision hardware products of space and land designation, for example, conical nose cones of rocket carriers, transient bays, ring ramps. The lamellar sheathing is made out of layers of the fibrous material impregnated with a polymeric binding. Its each layer represents an unrolling or a part of a cone unrolling made in the form of a sector of a ring or a sector of a circle with a central angle β. In the capacity of the unidirectional fibrous material use a fibrous material impregnated with a polymeric binding. The central angle β of a sector of a ring or a sector of a circle makes 12-360 degrees. Each of layers of the lamellar sheathing consists of the placed butted to each other sectors with the equal central angle γ making 1-30 degrees. In each sector of one layer the fibers of the unidirectional fibrous material are located at the similar for this layer angle φ to the central axis of the sector equal to 0- ±90 degrees. The butts of the sectors of each subsequent layer are shifted in respect to the butts of the sectors of the previous layer at an angle δ composing a part of the central angle γ of the sector for. The technical result of the invention realization consists in creation of the sheathing with the stable physical-mechanical properties.

EFFECT: the invention ensures creation of the sheathing with the stable physical-mechanical properties.

3 dwg

FIELD: spacecraft.

SUBSTANCE: proposed expulsion system with separation of liquid and gas for filling spacecraft at orbit contains pressurizing gas cylinders, pressure regulator, tank to be emptied with displacement diaphragm dividing the tank into liquid and gas spaces, transfer line with valve and tank to be with drain line. Propellant expansion compensator is installed additionally on transfer line between tank to be emptied and valve. Said compensator is made in form of reservoir with flexible separator of liquid and gas whose gas space is connected with gas space of tank to be emptied, and liquid space, with liquid space of said tank.

EFFECT: improved reliability of expulsion system, possibility of its operation after interruptions in spacecraft filling.

4 cl, 2 dwg

FIELD: space engineering; spacecraft flying in geostationary or high-altitude elliptical orbits.

SUBSTANCE: proposed spacecraft has module case with projecting members. Two opposite faces of each module perform function of radiators with built-in thermal tubes. Arranged in modules are engine unit and some heat-loaded units and onboard devices (number n). Other units (number k) , for example metal-hydrogen storage batteries are secured to engine unit and are heat-insulated from first units. Units and devices are secured to engine unit by means of brackets through heat-insulating gaskets. Unit is made in form of three-layer honeycomb panel where thermal tubes with heaters are laid. Each of k-units has thermal contact with axial U-shaped thermal tube embracing them. This thermal tube is brought in contact with evaporator of loop thermal tube connected with radiator by means of vapor line which is communicated with loop thermal tube and its evaporator through condensate lines. Radiators are mounted beyond boundaries of projecting parts shading zones on side of extravehicular space.

EFFECT: increased cooling effect of spacecraft temperature control system; reduction of mass of this system.

2 cl, 4 dwg

FIELD: space engineering; spacecraft flying in geostationary or high-altitude elliptical orbits.

SUBSTANCE: proposed spacecraft has module case with projecting members. Two opposite faces of each module perform function of radiators with built-in thermal tubes. Arranged in modules are engine unit and some heat-loaded units and onboard devices (number n). Other units (number k) , for example metal-hydrogen storage batteries are secured to engine unit and are heat-insulated from first units. Units and devices are secured to engine unit by means of brackets through heat-insulating gaskets. Unit is made in form of three-layer honeycomb panel where thermal tubes with heaters are laid. Each of k-units has thermal contact with axial U-shaped thermal tube embracing them. This thermal tube is brought in contact with evaporator of loop thermal tube connected with radiator by means of vapor line which is communicated with loop thermal tube and its evaporator through condensate lines. Radiators are mounted beyond boundaries of projecting parts shading zones on side of extravehicular space.

EFFECT: increased cooling effect of spacecraft temperature control system; reduction of mass of this system.

2 cl, 4 dwg

FIELD: space engineering.

SUBSTANCE: rotation apparatus comprises casing (10), shaft (30) set in bearings (11) inside the casing, and motor (13) for rotating the shaft. Shaft (30) is provided with radial cantilevers (281) whose ends have a number of boxes (282) with experimental objects. The apparatus is provided with devices for damping vibration of the casing. Each of the devices has spring mechanism (255) that flexibly secures the casing to member (280) of the unmovable bearing and electromagnetic mechanism for setting. The latter comprises conductor (254) secured to member (280) and exciting coil (253) that embraces it and is mounted on the surface of casing (10). The coil is connected with the spring mechanism so that casing (10) is held by electromagnetic force that effects on conductor (254) when the coil is excited. The apparatus may be provided with a pickup that measures the space between casing (10) and member (280) of the bearing and control unit that receives the signals from the pickup. When the space exceeds a given value, the unit controls the current in exciting coil (253) to reduce casing vibration.

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

FIELD: spacecraft.

SUBSTANCE: proposed method includes topping up of propellant tank with pressurizing gas and transfer of propellant components from emptied tank into receiving tank. Amount of propellant transferred from emptied tank is determined by ratio depending on pressurizing gas temperature at beginning and end of propellant transfer and temperature and pressure in gas space of emptied propellant tank at end of process.

EFFECT: improved reliability of spacecraft topping up system, simplified design of propellant tank and reliability of its sealing.

1 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

FIELD: space orbital stations.

SUBSTANCE: proposed system contacts autonomous subsystems located on cargo spacecraft and on space orbital station and including containers for liquid products with pipelines and valves. Container for liquid products of autonomous subsystem arranged on cargo spacecraft is made in form of bag of soft elastic cloth placed in rigid envelope and connected by flexible pipeline through hydraulic connector with change-over valve or through hydraulic connector with change-over valve and pump to similar container of other autonomous subsystem located on space orbital station. Space formed between bag and rigid envelope of autonomous subsystem located on cargo spacecraft communicates through valve with atmosphere of cargo spacecraft. Similar of container of autonomous subsystem located on space orbital station communicates through valve with atmosphere of space orbital station. Space orbital station is furnished with liquid wastes collecting system containing reservoir to be filled with liquid wastes in which space of bag is connected to semiconnector by means of flexible pipeline. Space formed between bag and rigid envelope communicates with atmosphere of space orbital station though valve or through valve and compressor. Invention makes it possible to create liquid products replenishment system providing utilization of liquid wastes from reservoir of autonomous subsystem at space orbital station by increasing efficiency of use of water container of autonomous subsystem of cargo spacecraft.

EFFECT: provision f utilization of liquid wastes.

3 cl, 2 dwg

FIELD: spacecraft equipment; control of rotation parameters of rotator with experimental objects and measurement of mass of these objects.

SUBSTANCE: proposed device has body with grooves (10a, 10b) and rotating shaft (30) whose both ends are fitted in bearings (11, 12) of grooves (10a, 10b). Lower end of shaft (30) is connected with engine (13). Four arms (24-27) are mounted horizontally and are secured on shaft (30) at one end; experimental boxes (20-23) are fitted on other ends. Objects of constant or increasing weight, such as plants or animals and human being are placed in boxes (20-23) where artificial micro-gravitation is formed by rotation for conducting experiments or work in space. Vibration of shaft (30) and boxes (20-23) is absorbed by bearings (11,12). Provision may be also made for special-purpose vibration generator. Side plates (1a-1d), acceleration sensors (2a-2d) and sensors (3a-3d) showing the distance between objects and plates are mounted in boxes (20-23). Mass of object is determined by means of computer unit on basis of signals received from sensors (2,3) at collision of object with side plate. These and similar signals from additional sensors are used for control of said vibration generator and motion of counter-weights (not shown) for elimination of unbalance of rotator and suppression of vibrations.

EFFECT: enhanced efficiency; simplified procedure.

42 cl, 64 dwg

FIELD: spacecraft equipment; control of rotation parameters of rotator with experimental objects and measurement of mass of these objects.

SUBSTANCE: proposed device has body with grooves (10a, 10b) and rotating shaft (30) whose both ends are fitted in bearings (11, 12) of grooves (10a, 10b). Lower end of shaft (30) is connected with engine (13). Four arms (24-27) are mounted horizontally and are secured on shaft (30) at one end; experimental boxes (20-23) are fitted on other ends. Objects of constant or increasing weight, such as plants or animals and human being are placed in boxes (20-23) where artificial micro-gravitation is formed by rotation for conducting experiments or work in space. Vibration of shaft (30) and boxes (20-23) is absorbed by bearings (11,12). Provision may be also made for special-purpose vibration generator. Side plates (1a-1d), acceleration sensors (2a-2d) and sensors (3a-3d) showing the distance between objects and plates are mounted in boxes (20-23). Mass of object is determined by means of computer unit on basis of signals received from sensors (2,3) at collision of object with side plate. These and similar signals from additional sensors are used for control of said vibration generator and motion of counter-weights (not shown) for elimination of unbalance of rotator and suppression of vibrations.

EFFECT: enhanced efficiency; simplified procedure.

42 cl, 64 dwg

FIELD: space technology.

SUBSTANCE: unconfined space of gas chamber of hydro-pneumatic compensator is subject to periodical change at the same average-mass temperature of heat-transfer agent. The ratio of Vi≤(Vi+l+nϕ) 1) is used to judge if leak-proofness corresponds to standard value, where Vi is volume of gas chamber of hydro-pneumatic compensator for i-th measurement, Vi+l is volume of gas chamber of hydropneumatic compensator for subsequent measurement, n is time interval between i-th and i+1 measurement, ϕ is standard value of volumetric loss of heat-transfer agent during specific time interval. Difference in unconfined spaces achieved between (i+1)-th and i-th measurement is used to determine real leakage of heat-transfer agent from system during specific time interval. Current value of unconfined space of system hydro-pneumatic compensator gas chamber is measured instead of measuring working pressure of the system for the same average-mass temperature of heat-transfer agent. Difference between measured spaces related to time interval between measurements has to be value of real leakage of heat-transfer agent observed during specific time interval.

EFFECT: simplified and reliable method of inspection.

FIELD: space engineering; servicing orbital stations, type MIR in space.

SUBSTANCE: proposed compartment has housing and oxidizer and fuel tanks secured on its frame and provided with fittings and supercharging system. Tanks with different components are combined in propellant modules which are diametrically opposite relative to each other. Compartment includes additionally oxygen bottles and water reservoirs mounted on said frame between propellant modules diametrically opposite relative to each other. Provision is made for electric heaters made from carbon cloth and located mainly in zones of water reservoirs. Outer surfaces of fuel and oxidizer tanks, oxygen bottles and water reservoirs and inner surface of housing are covered with temperature control coats. Outer surface of housing is coated with heat insulation.

EFFECT: simplified centering of compartment due to effective thermostatting of units located in this compartment.

2 dwg

FIELD: rocketry and space engineering; cryogenic stages of space rockets.

SUBSTANCE: according for first version, oxidizer supply unit is shifted in transversal direction and is secured in lower point of convex part of lower head plate of oxidizer tank, thus forming additional space in inter-tank compartment in axial direction; this space is used for displacement of cruise engine together with fuel tank towards oxidizer tank. According to second version, oxidizer supply unit is secured on concave part of lower head plate of oxidizer tank. Full suction of oxidizer from tank is performed by means of passages of intake unit introduced into concave part of lower head plate of oxidizer tank and used for coupling the lower zone of oxidizer tank with oxidizer supply unit inlet.

EFFECT: improved mass characteristics due to reduction of overall dimensions in length.

2 dwg

FIELD: rocketry and space engineering; cryogenic stages of space rockets.

SUBSTANCE: according for first version, oxidizer supply unit is shifted in transversal direction and is secured in lower point of convex part of lower head plate of oxidizer tank, thus forming additional space in inter-tank compartment in axial direction; this space is used for displacement of cruise engine together with fuel tank towards oxidizer tank. According to second version, oxidizer supply unit is secured on concave part of lower head plate of oxidizer tank. Full suction of oxidizer from tank is performed by means of passages of intake unit introduced into concave part of lower head plate of oxidizer tank and used for coupling the lower zone of oxidizer tank with oxidizer supply unit inlet.

EFFECT: improved mass characteristics due to reduction of overall dimensions in length.

2 dwg

FIELD: rocketry and space engineering; cryogenic stages of space rockets.

SUBSTANCE: according for first version, oxidizer supply unit is shifted in transversal direction and is secured in lower point of convex part of lower head plate of oxidizer tank, thus forming additional space in inter-tank compartment in axial direction; this space is used for displacement of cruise engine together with fuel tank towards oxidizer tank. According to second version, oxidizer supply unit is secured on concave part of lower head plate of oxidizer tank. Full suction of oxidizer from tank is performed by means of passages of intake unit introduced into concave part of lower head plate of oxidizer tank and used for coupling the lower zone of oxidizer tank with oxidizer supply unit inlet.

EFFECT: improved mass characteristics due to reduction of overall dimensions in length.

2 dwg

FIELD: spacecraft temperature control systems; removal of low-potential heat from on-board systems of spacecraft.

SUBSTANCE: proposed trickling cooler-radiator includes heat-transfer agent storage and delivery system, drop generator with acoustic oscillation exciting element, drop collector, transfer pumps and pipe lines. Trickling cooler-radiator is provided with heat stabilization system including heaters mounted on structural members of cooler-radiator and thermostatting units made in form of shield-vacuum insulation of these members. Said system is also provided with bypass pipe line laid between drop generator and collector and provided with volumetric expansion compensator (with electric heater) and automatic temperature control unit ensuring operation of heaters by signals from respective sensors. To reduce emission of heat-transfer agent, trickling cooler-radiator is provided with hydraulic accumulators at drop generator inlet and at drop accumulator outlet. Passages of output grid of drop generator have geometric and hydraulic characteristics varying from axis of symmetry towards periphery for smooth distribution of temperature field. Drop collector may be passive with inner surface formed by walls of one or several slotted passages through which heat-transfer agent is delivered for forming moving film.

EFFECT: enhanced efficiency and reliability.

2 cl, 2 dwg

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