Low-thrust cryogenic propulsion module
FIELD: classic and return launch vehicles.
SUBSTANCE: proposed low-thrust cryogenic propulsion module contains main cryogenic engine 10, two auxiliary engine 21, 22 to control position in space, cryogenic ergol supply tanks 31, 32, 33, 34, device for periodically building pressure in tanks 31, 32, 33, 34 and device to generate explosive pulses of main cryogenic engine at pulse mode during period of pressure building in tanks 31, 32, 33, 34. Device to periodically build pressure in tanks 31, 32, 33, 34 has heat exchange system connected with heat accumulator 61, 62 and device 71,72 to excite circulation of preliminary set amount of ergol in heat exchange system. Module contains additionally device for heating heat accumulator 61, 62 in period between two sequential explosive pulses.
EFFECT: improved mass-and-dimension characteristics of module, reduced time taken for execution of task.
17 cl, 14 dwg
The technical field to which the invention relates.
The present invention relates to cryogenic propulsion module low-thrust for use in the classical or return booster.
More specifically, the invention relates to cryogenic propulsion module low-thrust, allowing the transfer of a satellite from one orbit to another, while the motor module can be embedded in the satellite or to provide a detachable step.
The level of technology
In the field of launching satellites the main direction of development is to improve the ratio between launch into geostationary orbit weight and weight of a particular booster.
The classic way to start is to launch of the satellite into transfer orbit, called gap (geostationary transition orbit), and then transfer the satellite to geostationary orbit using an include at the top of a jet engine on a two-component liquid fuel, with this apogee motor module is built into the satellite.
It was also proposed to transfer a satellite from a low orbit to geostationary orbit using solar thermal stage, which uses liquid hydrogen.
This method is described, for example, in the article, J.A. Bonometti, C.W. Hawk "Solar thermal rocket research apparatus and proposed testing" (University of Alabama - 1994).
Tacos the way mentioned in article J.M. Shoji "Potential of advanced thermal propulsion. Orbit raising and maneuvering propulsion: research status and needs", published in Progress in Astronautics and Aeronautics, AIAA - Volume 87, p.30-47.
According to this known method, which is illustrated in Fig. 2, the sunlight is concentrated by a parabolic mirror 5 in a solar oven 6 in which the hydrogen is heated to a temperature of about 2000 K. Then reduce the hydrogen pressure in the nozzle of the engine 8, providing improved ejection velocity (7500-8000 m/s), i.e. specific tractive pulse of about 750-800 C. In Fig. 2 shows a diagram of such a system in which the satellite 2 is connected with one side of the booster through the connecting device 1 and the other side with the hydrogen tank 3 through design 4. Position 7 schematically indicated device receiving liquid hydrogen to power a solar oven 6 engine 8.
This device, which was never used in practice, in theory it should give the possibility to increase the weight displayed on the geostationary orbit. However, this configuration has many flaws.
In particular, to achieve a temperature of 2000 To the need to ensure that the concentration ratio of the solar flux from 5000 to 8000. This suggests a mirror of very high quality, which is very difficult to obtain in terms of restrictions on vehicle weight. In addition, guidance on canadaline to be very accurate, order ± 5' on two axes, which raises difficult problems of control orientation.
The dimensions of the tank 3 for liquid hydrogen also create difficulties. Indeed, to obtain, for example, the total traction pulse value 30 MNS, you must use a tank containing 4000 kg of liquid hydrogen, which corresponds to a capacity of 60 m3(this means, for example, the diameter of 4.2 m and a height of 5 m).
Because of these disadvantages, the development of translation systems on orbit based on solar thermal stage using liquid hydrogen remains hypothetical.
Another technology that gives you the ability to increase the mass delivered to geostationary orbit or gap, suggests the booster with the latest cryogenic stage, which allows the use of tanks relatively reduced volume with easier placement into the booster. Thus, to obtain General traction pulse value 30 MTL cryogenic stage with liquid hydrogen and liquid oxygen requires masses of Argos 6600 kg, but the capacity of the tanks is only 22 m3.
However, the currently used cryogenic stages require the use of pumps, which increases their value.
Some authors have proposed to perform cryogenic stage with food under pressure without the use of pumps, however, these ideas did not find spiral is Noah implementation. And indeed, in practice, the hydrogen pressure should always be higher oxygen pressure, to provide regenerative cooling of the combustion chamber. From this it follows that to create the pressure required is too large mass of helium.
The problem to which the present invention is directed, is to eliminate these obstacles, in particular in ensuring the translation of satellites in different orbit using simpler, less heavy and bulky device in comparison with the known devices, allowing to avoid the use of pumps and enabling the use of engines and tanks of moderate size for Argos in order to reduce the dimensions of the step motor, is required to transfer the satellite to another orbit. The term "argol"as it is known, understood as rocket fuel or a component of rocket fuel.
In accordance with the invention the solution of this problem is achieved by creating cryogenic propulsion module low-thrust, providing thrust in the range from 100 to 1000 N. The motor module according to the invention is characterized in that it contains at least one main cryogenic engine having a pressure in the combustion chamber from 2102up to 103kPa, at least two auxiliary the orientation of the engine, at least first and second tanks power cryogenic elalami, means for creating pressure pulses in these tanks supply and launcher main cryogenic engine in a pulsed mode during pulse pressure in these tanks power. The duration of the period between two successive explosive pulses of the engine is approximately 1 hour and 30 minutes to 12 hours. This means the periodic creation of pressure in these tanks power contain at least one heat exchange system connected to a heat accumulator, and means to circulate a predefined number of Argos in the specified heat exchange system. The module further comprises also means for heating the heat accumulator during periods between two successive explosive impulses.
Associated with the tank erholen heat accumulator can be heated at least partly by using the solar thermal system, such as flat solar thermal installation, which is the ratio of the absorption capacity for emissivity (α/ε) is greater than 1 and is provided with spericolata on its rear surface.
Heat accumulator can also be heated at least partially through the recovery of heat losses is toplivnogo battery, functioning on the evaporated argosh.
The fuel cell power supply can be powered cold pairs of Argos leaving the heat exchanger, designed to maintain a constant temperature in the zone of selection in the tank with erholen.
Heat accumulator can also be heated at least partly by electrical heating.
The accumulation of heat within the heat accumulator in the best case scenario is carried out with material, performing a phase transition, such as alkali metal or a hydrocarbon.
According to a particular exemplary embodiment of the propulsion module contains the first and second tanks with erholen to power the main engine, while argali together evaporate in the associated tanks with thermal batteries are thus to ensure a constant composition of the mixture.
In accordance with the special advantages of the exemplary embodiment of the cryogenic propulsion module contains the first and second main tanks with elalami and at least first and second auxiliary tanks with elalami, which form the buffer tanks. The auxiliary tanks are designed to create pressure by a specified means of creating pressure and have such dimensions that allow for orbital maneuver through periodic power main engine and a full is emptied at the end of the explosive impulse. Thus means are provided for refilling these auxiliary tanks from the respective main tanks between two successive explosive pulses, and the pressure in the main tank is maintained below the pressure supply main engine
In this case, according to particular exemplary embodiment an auxiliary tank covered with insulation and installed inside the main tank.
List of figures
Other features and advantages of the present invention will become apparent from the following description of some embodiments of the invention, given as examples, with reference to the accompanying drawings, on which:
figure 1 depicts in schematic General view of a cryogenic propulsion module in one of the variants of its implementation and the satellite associated with the module,
figure 2 schematically depicts a solar thermal stage in accordance with the prior art and the satellite associated with this solar thermal stage,
figure 3 schematically presents a set of basic functional bodies in the exemplary embodiment of the cryogenic propulsion module with solar heating,
4 and 5 depict, respectively, on the form in section, and the front is an example running flat solar thermal plant with integrated thermal battery, Teploobmennik is, applicable device according to the invention,
6 schematically depicts in sectional view the connection between the flat solar thermal installation, and a separate thermal battery
7 in schematic image shows an example of the solar thermal installation associated with flat reflecting mirrors and applicable in the device according to the invention,
Fig depicts in schematic form an example of a complete solar thermal system associated with a parabolic reflecting mirrors,
figure 9 schematically presents a set of basic functional bodies in one example implementation cryogenic propulsion module using auxiliary buffer tanks,
figure 10 is a schematic representation of a set of basic functional bodies in another example, execution of cryogenic propulsion module using auxiliary buffer tanks and preliminary evaporation of Argos,
figure 11 is a schematic representation of a set of basic functional bodies in another exemplary embodiment of the cryogenic propulsion module using a fuel cell,
Fig illustrates the inclusion of a heat exchanger in the bottom of the main tank with erholen applicable in cryogenic propulsion module according to the invention,
on Fig site presents the heat exchanger is, shown in the example of execution according pig,
Fig depicts a partial view of the main tank with erholen, which can be used in cryogenic propulsion module according to the invention and contains a buffer tank inside the main tank.
Information confirming the possibility of carrying out the invention
In Fig. 1 presents in a schematic General view of one exemplary embodiment of the cryogenic propulsion module 100 in accordance with the invention. The propulsion module 100 contains the main engine 10 with fuel type oxygen-hydrogen, the pressure in the combustion chamber which is approximately 2·102up to 103kPa and is low enough to create a wall heat flux at five to ten times lower than in conventional cryogenic engine. This allows you to use most of the engine 10 is a simplified regenerative cooling or even cooling through radiation or through film.
The main engine 10 may be a single engine mounted in a universal joint, or it can represent a set of at least three main cryogenic engines, individual rod which is adjustable by means of an adjustable pressure relief in the supply systems of Argos.
The main engine 10 or the set of basic motors has little traction on adca from 100 to 1000 H, to reduce the size and to reduce the dimensions stage propulsion system. The use of multiple main thrusters allows, for example, to reduce the length of a degree more than 3 m in comparison with the known structures.
Cryogenic propulsion module 100 may contain from two to six orientational engines, such as engine 21 control yaw and the motor 22 to control the roll.
The supply of cryogenic propulsion module 100 can be set on the module solar photovoltaic panels 51, but may also be performed using detachable connectors associated with the satellite 200, intended for placing in orbit using cryogenic propulsion module 100. In another embodiment, the power supply may be provided with a fuel cell power supply, such as fuel element 270 in the example of execution according to Fig. 11, which can be powered by the evaporation of the cryogenic Argos.
The power of the main engine 10 is carried out in a pulsed mode by creating pressure in the main tanks 31, 32, are stored, respectively, liquid hydrogen and liquid oxygen. Since the operating pressure is low, the coefficient of the structural strength of the tanks 31, 32 remains within reasonable limits.
The tanks 31, 32 p. the villas of cryogenic Angola, such as H2O2the pressure is created without the use of pumps, by simple evaporation of a predetermined amount of each of Argos in thermal battery, such as, for example, the battery 60 in Fig. 4 and 5 or the battery 160 in Fig. 6. Heat accumulator 60 is connected with a heat exchange system 70 and electric micronesica to circulate a certain number of Argos in the heat exchange system 70.
In Fig. 1 shows an example of Micronase 71 associated with the first tank 31 and thermal accumulator 61.
Heat accumulator 60, 160 is heated during the periods between two successive explosive pulses of the main motor 10.
Heat accumulator can be heated either by solar or electric heat, or by recovery of heat losses of the fuel cell, which operates with the help of evaporated Angola, or using combinations of these three methods.
Heat accumulator is heated in the interval between two successive explosive pulses of the engine or of the engine 10, while the duration of this interval is 1 hour 30 minutes to 12 hours depending on the eccentricity of the orbit. The number of consecutive explosive pulses of the main motor 10 may be, for example, from 10 to 30. These pulses is Sy performed at perigee or apogee of the orbit, so orbital maneuver is made in discrete increments, given the deliberately low thrust cryogenic propulsion module 100. However, the time periods between two successive explosive pulses are not idle periods, and are used for heating the heat accumulator.
In Fig. 4 and 5 shows an example of performing heat accumulator 60 flat solar thermal installation 60A, which directly heats the heat accumulator 60 by contact with tubes circulation of hydrogen heat exchange system 70. Layer 60b of superisolated is located on the rear side of the heat accumulator 60 and the heat exchanger 70.
A heat accumulator 60 in optimum performance it is made of a material making the phase transition, such as alkali metal or a hydrocarbon, which helps to reduce weight.
Flat solar thermal installation 60A may be provided with a coating with a controlled emissivity (the ratio of the absorption capacity/emissivity α/ε≫1), so that the solar thermal installation can achieve a sun-resistant temperature above 100°C. Solar thermal installation according to Fig. 4 and 5 allows the angles of incidence of the solar flux in a wide range of angles.
Alternatively, in Fig. 6 shows an example of a run-flat Solna is Noah thermal installation 160A, which is provided with a layer 160b of superisolated on the rear side and is used to heat a separate heat accumulator 160 through the system of tubes 166 on the front surface of the solar thermal installation 160A and related microventilation 170b system a circulation of fluid.
Unlike solar thermal actuator using flat solar thermal installation 60A avoids the need for precise aiming - error procedure ±20° on two axes is quite valid.
Surface, and hence the mass of the solar thermal installation 60A can be determined by using the flat concentrating mirrors 161 (Fig. 7) or a parabolic cylindrical mirrors 172 (Fig. 8) without increasing demands on the accuracy of pointing at the Sun.
Figure 3 presents a schematic diagram of a pressure in the tanks 31, 32 through the heat storage with heat accumulators 61, 62, which are connected with solar thermal installations and main tanks 31, 32 of liquid hydrogen and liquid oxygen.
In the example of execution according to Fig. 3 electronic control circuit, powered by a solar panel 51, provides power to an electric Micronesia 71, 72, associated with the tanks 31, 32. The solar panel 51 can be mounted on the motor unit on the satellite, intended for the of Yoda into orbit. The electrical connection between the solar panel 51 and an electric control circuit can be provided by means of the coupling device with an extension cord. The propulsion module 100 may also be made integral with the satellite.
Team electric Micronase 71, 72 inject liquid argali in thermal batteries 61, 62, which causes the temperature rise of Angola to approximately ambient temperature and due to their feed lines 105, 106 increases the air pressure in the respective tanks 31, 32.
Upon reaching the target pressure Micronase 71, 72 stop. Management Micronase 71, 72 is provided with electronic circuit 110 controls that are associated with the sensors 101, 102 pressure, measuring the pressure in the tanks 31, 32.
After creating pressure in the tanks 31, 32 is sufficient to open the solenoid valves 91, 92 power elalami main engine 10 and then to produce ignition in the engine 10 by the electric discharge to ensure the orbital maneuver.
Presented on Fig. 3, the system can be improved through the use of auxiliary buffer tanks 33, 34 in combination with the main tanks 31, 32 for Angola, as shown in Fig. 9-11. For the purpose of greater clarity in these drawings are not submitted to the electronic circuit 110 controls, sensors 101, 102 and pressure source 51 elektricheska the power included in the system, but have already been described with reference to the previous example implementation.
In those cases, when the mass of Angola spent during one maneuver, is of the order of 100 kg, can be used to maneuver the buffer tanks 33, 34 small size, in which it is easier to create pressure than in the main tanks 31, 32. This allows, in particular, to reduce the coefficient of the structural strength of the main tanks 31, 32, as they are subjected to a moderate pressure of about 102kPa.
As is clear from Fig. 9, the auxiliary tanks 33, 34, which in the initial state are under low pressure, after explosive pulse can be filled with erholen with its submission of the main tanks 31, 32 by opening the valves 93, 94, installed between the main tanks 31, 32 and auxiliary tanks 33,34.
Then in the auxiliary tanks 33, 34 is pressurized by Micronesian 71, 72, which inject liquid argol in thermal batteries 61, 62. The valves 103, 104 in the pipes 105, 106 pressure in the main tanks 31, 32 remain closed.
Thus, the pressure in the auxiliary tanks 33, 34 may be increased, for example, from 102bar to 5102kPa. Upon reaching the set pressure opens the valve 91, 92 power of the main engine 10 and may be orbital maneuver.
During the explosive impulse of the main engine 10, the pressure in the buffer tank 33, 34 is maintained approximately constant by the controlled actuation of Micronesia.
The gas flow circulating in thermal battery 61, 62, enables the operation is also the orientation of the motors 21, 22. The power of these engines with gaseous erholen is along the lines 121, 122 through the valves 107, 108, which are installed between the heat accumulators 61, 62 and buffer tanks 33, 34.
It should be noted that in the examples according to Fig. 3 and 9, the main motor 10 is powered by liquid elalami.
It may be desirable to apply to the engine 10 argali in the vaporized state, in order to avoid difficulties freezing or large fluctuations in the composition of the mixture that can be created, given the small size of the main engine 10.
In Fig. 10 shows an example where not only the auxiliary engines 21, 22, 23 of the feed gas, but the main engine 10 is fed with gaseous elalami, evaporated in thermal batteries 61, 62.
In the example of execution according to Fig. 10 argali submitted by Micronase 71, 72 in thermal batteries 61, 62, evaporate. The heat capacity of the batteries should be selected accordingly. Evaporated argali after passing through thermal battery 61, 62 are not sent back to the buffer tanks 33, 34, and injected directly into the main on the " 10 and auxiliary engines 21-23.
As an example, you can specify that for evaporation by heating 20 kg of liquid hydrogen requires 37 MJ of energy, which implies that the average power 2500 watts for 3 hours. Solar thermal installation area of 2.2 m2it is quite sufficient to ensure such capacity.
It should be noted that the filling of the buffer tanks can be carried out by microgravity. For this purpose it is necessary to ensure that liquid argol was always present on the side of the selection from the main tank.
To resolve this problem, ensure that the slow circulation of liquid Argos along the walls of the tank and produce local cooling of selected Argos exchanger.
An example of such a device called ATVE (Active Thermodynamic Vent System - active thermodynamic selection system) described in the publication E.C. Cady and A. OIsen "Thermal Upper Stage Technology Demonstration Program", AIAA 96-3011-32nd AIAA Joint Propulsion Conference, Lake Buena Vista, July 1996.
In Fig. 12 and 13 shows an example implementation of such a device in the main tank 31 of the motor module according to the invention.
The bottom of the tank 31 is occupied by the selected volume of Argos cooled via a heat exchanger 370, which may contain a heat-exchange tube 375 with ribs. A small electric pump 371 circulate chilled fluid in the tank through the Central tube 380. Control valve 390 pozvolyaet the fluid, and cold pairs then are taken out of the tank through the tube 391. Selection of cold vapor is produced only during the explosive impulse of the main engine, which is fed through the lower pipe 341.
Figure 11 shows an example run using fuel cell 270 power, which can continuously eat elalami from the main tanks 31, 32 and auxiliary tanks 33, 34. In particular, the fuel cell 270 may eat cold pairs of Angola leaving the heat exchanger, supporting a constant temperature selected volume in each main tank.
Through the power bus 280 fuel cell 270 power supply can supply the power needed for different consumers (pumps, valves, electric heaters, and unit equipment motor module (Central inertial unit, onboard computer, radio).
The efficiency of the fuel cell 270 power is approximately 50%, while losses are used for heating all or part supplied to the main motor 10 Angola module 260, which consists of a heat exchanger and a heat accumulator. Thus, dissipation of the fuel element 270 power supply power stored in the battery to heat the liquid argola that allows you to create pressure in the auxiliary tanks 33, 34 and to feed the main engine 10 and orientation engines 21-23.
If the fuel cell 270 power develops an average power of 1 kW, accumulated over three hours energy reaches approximately 11 MJ.
In that case, when using the embodiment of Fig. 11S fuel element 270 power instead of solar thermal system 60A as the heat source, the management orientation in space by means of auxiliary engines 21-23 can be carried out regardless of orientation to the Sun, which gives more freedom of positioning.
In Fig. 14 shows an embodiment in which the buffer tank 133 is located inside the main tank 131. This allows you to make the design more compact and to reduce thermal losses.
Buffer tank 133 provided with an outer layer 139 insulation and connected with the valve 193 filling line 138 creating pressure in the tank.
Buffer tank 133 is mounted above the heat exchanger 370, which may be similar to the heat exchanger in Fig. 13. The presence of the buffer tank 133 embedded in the main tank 131, helps retain fluid near the heat exchanger 370, especially at the final stage of the operation. Tube 380 circulation is shifted to the side of the main tank 131.
In the motor module according to the invention, the thrust vector control can be performed in three different ways, with the stabilization feature on the corner of CR is provided on, at least by means of two auxiliary motors 21, 22.
According to the first embodiment, the main motor 10 is connected with two pairs of auxiliary engines control in pitch and yaw relative to an axis parallel to the main engine 10, which is fastened in a fixed position.
According to the second variant the main engine 10 is installed in the universal joint. Two Electromechanical cylinder provide its orientation relative to the motor module.
According to a third variant of the function of the main engine, there are three or four engines, the thrust of which can be adjusted using a proportional valve, which distributes the flow of Argos. This allows you to control the position of the thrust vector relative to the center of gravity.
In the case when the cryogenic module 100 of the engine according to the invention uses solar thermal installation, it should be noted that the requirements of the guidance provided is very easy (tolerance 20° on two axes), while the use of solar thermal solutions technology provides guidance with accuracy within 5' along the two axes.
Thus, cryogenic propulsion module low-thrust has a lower volume due to the fact that the average density of Angola is 0.3 compared to dps what testu 0,07 solutions solar thermal technology. This facilitates the layout of all components in one casing. Furthermore, the reduced dry weight of the module according to the invention, since the total weight of thermal equipment is smaller and does not exceed 10% by weight tanks (compared to approximately double the magnitude of this ratio in the solutions solar thermal technology). In addition, there is no need primary hub of solar energy. And, finally, the motor module according to the invention allows to reduce the total duration of the assignment.
When compared with known cryogenic upper stage, corresponding to the prior art, the motor module according to the invention with similar technology provides a reduction in dry matter due to the fact that it does not use any helium environment or plate for reducing the pressure of the gas, and also due to the fact that the basic engine has less weight and much smaller. In addition, the main engine has a lower cost, as well as other components of smaller size.
It should be noted that in the solutions of cryogenic last stage powered by creating pressure in the tank in a known manner, the low pressure in the combustion chamber in combination with higher thrust leads to the fact that the engine has very large dimensions, but does not give a lower ratio in the cross section compared to the jet, the m turbo pump, that is, a lower specific traction and momentum.
In contrast, the solution proposed in the framework of the present invention makes it possible to combine low pressure in the combustion chamber with small dimensions due to perform many apsidal explosive impulses.
Low pressure in the combustion chamber reduces heat flow, which is approximately eight times lower than in the engine with the turbo pump, which allows the use of simplified regenerative cooling or even cooling by radiation.
In addition, the use of solar heating at least a pressure in the tanks eliminates the need for additional weight and relieves constraints associated with the compression of helium.
Finally, the invention allows to coordinate the sequence of intermittent explosive pulses with phases of the heating heat accumulators.
1. Cryogenic propulsion module low-thrust, providing thrust from 100 to 1000 N, characterized in that it contains at least one main cryogenic engine (10)having a pressure in the combustion chamber 2·102up to 103kPa, at least two auxiliary orientation of the engine (21,22)at least first and second tanks (31, 32, 33, 34) for feeding the cryogenic elalami, means(110, 71, 72, 61, 62) periodic build up pressure in the criminal code of the connected tanks (31, 32, 33, 34) and power tools (110, 91, 92) launch explosive pulses of the main cryogenic engine (10) in the pulse mode during the periodic creation of pressure in these tanks (31, 32, 33, 34) of the food, and the duration of the period between two successive explosive pulses is approximately 1 hour and 30 minutes to 12 hours, means the periodic creation of pressure in these tanks (31, 32, 33, 34) the food contains at least one heat exchanger system (70, a)associated with thermal accumulator (60; 160; 260) and means (71, 72) circulate a predefined number of Argos in the specified heat exchange system (70, a), and the module further comprises means (161, 162, 270) for heating the heat accumulator (60; 160; 260) during the periods between two successive explosive impulses.
2. The motor module according to claim 1, characterized in that the heat accumulator (60; 160) is configured to heat at least partially using a solar thermal system (60A, 160A).
3. The motor module according to claim 2, characterized in that the heat accumulator (60; 160) is made with the possibility of heating by means of a flat solar thermal system (60A, 160A), which is the ratio of the absorption capacity for emissivity (α/ε) is greater than 1 and is provided with spericolata (60b; 160b) on svezhina surface.
4. The motor module according to claim 2 or 3, characterized in that the solar thermal installation (60A; 160A) associated with the system planar or parabolic cylindrical concentrating mirrors (161; 162).
5. The motor module according to any one of claims 1 to 4, characterized in that the heat accumulator (260) configured to heat at least partially through the recovery of heat losses of the fuel cell (270) power supply operating in the evaporated argosh.
6. The motor module according to any one of claims 1 to 5, characterized in that the accumulation of heat within the heat accumulator (60; 160; 260) is carried out with material, performing a phase transition, such as alkali metal or a hydrocarbon.
7. The motor module according to claim 5, characterized in that the fuel cell (270) eats cold pairs of Argos leaving the heat exchanger (37), designed to maintain a constant temperature in the tank (31) with erholen.
8. The motor module according to any one of claims 1 to 7, characterized in that the heat accumulator (60; 160; 260) is configured to heat at least partially by electrical heating.
9. The motor module according to any one of claims 1 to 8, characterized in that it contains at least a single main cryogenic engine (10), mounted in a universal joint.
10. Motor mod is l according to any one of claims 1 to 8, characterized in that it contains at least three main cryogenic engine, individual rod which is adjustable by means of an adjustable pressure relief in the supply systems of Argos.
11. The motor module according to any one of claims 1 to 8, characterized in that it contains the first and second tanks (31, 32) with erholen to power the main engine (10), while argali together evaporate related to tanks (31, 32) thermal battery (61, 62) thus to ensure a constant composition of the mixture.
12. The motor module according to any one of claims 1 to 11, characterized in that the tank (31, 32) of the power supply is equipped with a pump (371) circulation, mounted on the body of the associated heat exchanger (370)located on the side of the selection.
13. The motor module according to any one of claims 1 to 12, characterized in that it contains at least first and second main tanks (31, 32) with elalami and at least first and second auxiliary tanks (33, 34) with elalami, which form the buffer tanks and auxiliary tanks (33, 34) is arranged to create a pressure by a specified means of creating pressure and have dimensions allowing orbital maneuver through periodic power main engine (10), and completely emptied at the end of the explosive impulse, when this provides a means for refilling is shown an auxiliary tanks (33, 34) from the respective main tanks (31, 32) between two successive explosive pulses, and the pressure in the main tanks (31, 32) is maintained below the pressure supply main engine (10).
14. The motor module according to any one of claims 1 to 13, characterized in that at least one auxiliary tank (133) covered with insulation and installed inside the main tank (131).
15. The motor module according to any one of claims 1 to 14, characterized in that it contains a source of electrical power, formed fuel element (270) power, fed by evaporation of the cryogenic Angola.
16. The motor module according to any one of claims 1 to 14, characterized in that it contains a source of electrical power, well-educated, at least one solar panel (51)attached to this module.
17. The motor module according to any one of claims 1 to 14, characterized in that it contains a source of electrical power, formed detachable connector associated with the satellite (200), intended for placing in orbit with the help of this module (100).
FIELD: rocket and space engineering.
SUBSTANCE: proposed liquid-propellant rocket engine has chamber. Bypass main line with flow rate regulator passing part of fuel into chamber by-passing cooling duct is installed in parallel to line of fuel delivery for cooling.
EFFECT: increased service life of engine chamber owing to reduction of thermal stresses in inner wall.
FIELD: rocketry, in particular, liquid-propellant rocket engines using helium as a cooler of the engine chamber body.
SUBSTANCE: the liquid-propellant rocket engine has an engine chamber consisting of a combustion chamber and a nozzle, having regenerative-cooling ducts, turbopump assembly including centrifugal pumps of oxidizer, fuel and helium, neutral gas generator fed from the pumps of oxidizer and fuel, and the outlet of the helium pump is coupled to the regenerative cooling passage of the combustion chamber, whose outlet is coupled to the mentioned gas generator, the outlet of the gas generator is coupled to the turbine of the turbopump assembly, whose outlet is coupled to the oxidizer supply line to the combustion chamber mixing head. Besides, cooling of the chamber nozzle is effected by fuel, which, having passed through the regenerative cooling ducts, is supplied to the mixing head. The combustion chamber of the engine chamber and the gas generator operate at a stoichiometric relation of the fuel components. Introduction of the helium additive to the combustion products of the main fuel components to the neutral gas generator and further to the engine combustion chamber makes it possible to enhance the engine specific thrust pulse approximately by 20S, and, with regard to denial of screen cooling, approximately to 30S and more.
EFFECT: enhanced engine specific thrust pulse.
1 cl, 1 dwg
FIELD: liquid-propellant rocket engines.
SUBSTANCE: proposed system includes lines of different propellant components, gas lines and gas generator. Mounted in front of main turbo-pump unit with multi-stage propellant component pump and drive gas turbine is booster turbo-pump unit with propellant component pump and drive single-stage hydraulic turbine fed from "n" stage of main pump. Working passage of hydraulic turbine is located in connecting line between outlet of "n" stage and inlet of "(n+1)" stage of main pump.
EFFECT: simplified construction; reduced mass of liquid propellant.
3 cl, 1 dwg
FIELD: aero-space engineering; flights in atmosphere and in space.
SUBSTANCE: proposed flying vehicle has casing, gas shock absorber of combustible fuel with control unit, cylinder and piston, two exhaust pipes and two resilient safety stops for piston which is engageable with shock absorber and performs reciprocating motion inside cylinder. Gas shock absorber is communicated with control unit for metered delivery of fuel to shock absorber. Mounted inside cylinder is cylindrical strut rigidly connected with rear wall of cylinder and with two said safety stops. Piston is provided with projection inside strut and additional shock absorber of fuel being ignited is provided inside strut behind piston projection. Control unit may be used for control of two shock absorbers due to alternating metered delivery of fuel. Provision is also made for two bent exhaust pipes rigidly connected with cylindrical strut, exhaust nozzle behind gas shock absorber which is rigidly connected with rear wall of cylinder and two booster jet engines rigidly connected with casing. Additional gas shock absorber and exhaust nozzle perform functions of jet engine (which was earlier mounted on piston) and this engine may be excluded.
EFFECT: enhanced reliability and durability; reduced mass of flying vehicle.
FIELD: aero-space engineering; flights in atmosphere and space.
SUBSTANCE: proposed flying vehicle has casing, gas shock absorber of combustible fuel with control unit, cylinder and piston, two exhaust pipes and two resilient protective stops for piston. Piston is engageable with shock absorber and performs reciprocating motion inside cylinder. Gas shock absorber is connected with control unit for metered delivery of fuel. Cylindrical strut mounted inside cylinder is rigidly connected with its rear wall and with two protective stops. Piston is provided with projection located inside cylindrical strut. Bent exhaust pipes are also located in this strut. Provision is made for mechanical shock absorber which is connected with casing and is located inside cylinder in front of piston and exhaust nozzle mounted behind gas shock absorber and rigidly connected with rear wall of cylinder. Gas shock absorber and exhaust nozzle perform function of jet engine (which was mounted on piston in previous construction), therefore there is no need in jet engine.
EFFECT: simplified construction; reduced mass of flying vehicle.
FIELD: the invention refers to fuel systems primarily to transport spaceships that provide refueling of orbital stations of the "Mir" type.
SUBSTANCE: the proposed compartment has a ring frame, tanks with an oxidizing agent and fuel with armature and supercharging systems corresponding to these components. At that the compartment is fulfilled with two fuel modules each of which has a separate mounting frame with the armature and the supercharging systems installed on it. The fuel modules are located relatively to each other on transversally opposite sections of the ring frame with provision of diametrically opposite location of the tanks with equal components of fuel. The presence of individual mounting frames facilitates conduction of arrangements at assembling, testing and replacing of the modules.
EFFECT: creation of a compartment of refueling components simple in manufacturing and having increased reliability.
FIELD: means of loading by high-density gases mainly of capacities of power units of space vehicles.
SUBSTANCE: the offered method consists in maintaining of pressure in the tank being loaded at which xenon overflows from the filling tank to the tank being loaded. After balancing of pressures in the filling tank and in the tank being loaded topping-up is performed at a temperature below the critical one in the transfer conditions with monitoring of the topping dose. Prior to installation of the tank to be loaded on the space vehicle, liquefied xenon is brought to the gaseous state by heating. Two check valves are installed in succession in the offered device in the loading line, a pressure accumulator is positioned between these valves. The gas accumulator is made in the form of a vessel separated into two cavities by a membrane. One of the cavities is linked with the pressure source, and the other - with the loading line between the mentioned check valves.
EFFECT: enhanced quality of loading due to a more tight filling of the tank being loaded, preserved fuel in case of cancellation of space vehicle starting, which results in a reduced cost of starting.
7 cl, 1 dwg
FIELD: spacecraft engine plants, mainly geostationary communication satellites; orientation of satellite and correction of its orbit.
SUBSTANCE: proposed module of spacecraft engine plant includes low-thrust engine (or engine package) mounted in biaxial suspension on rod and components of engine plant pneumatic system. Engine is mounted on one end of rod and pneumatic system components are mounted on other end of rod. Articulation securing the rod to spacecraft case is located between engine (or engine package) and pneumatic system components. Biaxial suspension makes it possible to perform independent rotation of engine (or engine package) relative to rod at two perpendicular axes through 90° around each axis. Rod securing articulation makes it possible to rotate the rod through 360° relative to spacecraft case. Provision is made for shaping correcting pulse during operation of low-thrust engine (or engine package) in any direction for maintenance of point of sight of spacecraft on orbit and creating required control moment around center of mass of spacecraft for its orientation.
EFFECT: enhanced efficiency.
4 cl, 2 dwg
FIELD: construction of spacecraft and arrangement of engine plants on them.
SUBSTANCE: proposed spacecraft includes load-bearing cylindrical body of central compartment with radial sheathing panels connected to it. Mounted on "north" and "south" sides of spacecraft are thermal panels for payload instruments and spacecraft service systems. Solar batteries are mounted on their outsides. Connected to central compartment on "zenith" and "ground" sides of spacecraft are package panels. Connected to package and thermal panels and to sheathing panels on " west" and "east" sides of spacecraft are engine panels. Mounted symmetrically on engine panels are sections of engines of equal thrust. Each section includes electrical jet engine (traction engine, for example stationary plasma engine) and gas jet engine (for dampening spacecraft after separation and relief of flywheels). Lines of action of engine thrust are inclined at different angles relative to spacecraft body axes and are shifted relative to center of mass of spacecraft.
EFFECT: possibility of using proposed spacecraft with engine module as accompanying payload at twin injection into space of main payload.
5 cl, 7 dwg
FIELD: spacecraft propellant equipment; systems for topping-up orbital stations, type MIR.
SUBSTANCE: proposed module has housing, frame with oxidizer and fuel tanks secured on it and respective fittings and tank pressurization system. Modules are provided with two-chamber bellows. One chamber of bellows is connected with tank pressurization system and other chamber is communicated with liquid cavity of respective tank. Bellows are secured on frame at perpendicular orientation of longitudinal axes of symmetry relative to longitudinal axis of symmetry of propellant module coinciding with direction of flight.
EFFECT: enhanced reliability due to avoidance of additional loads on tanks caused by temperature variations.
FIELD: propellant systems for aero-space craft, reservoirs for storage of cryogenic propellant (hydrogen, for example).
SUBSTANCE: proposed reservoir has hermetic protective casing, heat-insulated inner reservoir, filling, venting, pressurizing and propellant supply pipe lines and intake unit. Inner reservoir is secured to protective casing on two supports containing thin-walled bodies of revolution, taper in shape with heat-protective circular spacers and shield-vacuum heat insulation. One thin-walled body of revolution is connected with bush which is telescopically connected in its turn with cylinder attached to inner reservoir, thus ensuring relative motion of inner reservoir and protective casing. It is preferable that intake unit is provided with shut-off valves rigidly interconnected together and communicated with cavities near respective bottoms of inner reservoir by means of propellant intake pipe lines. These valves are controlled by means of electro-pneumatic system.
EFFECT: possibility of protracted storage of cryogenic propellant without venting when transport facility is at rest; continuous suction of propellant at alternating acceleration during motion.
2 cl, 4 dwg
FIELD: spacecraft propellant equipment for topping-up systems of orbital stations, type MIR.
SUBSTANCE: proposed propellant module has body and frame, oxidizer and fuel tanks, respective fittings and pressurization system. Body is made from two thin-walled truncated cones with oxidizer and fuel tanks arranged inside them. Cones are provided with electric heaters in form of jackets made from carbon cloth and secured on inner surfaces of cones. Circular clearances are formed between electric heaters and tanks. Fittings and tank pressurization systems are secured on outer surfaces of cones which are provided with detachable joints with frame. Frame is provided with holes for circulation of air through said clearances.
EFFECT: enhanced efficiency of thermostatic control of oxidizer and fuel tanks.
FIELD: spacecraft engine systems; construction of solar sails.
SUBSTANCE: proposed craft has hull, main and additional circular reflecting surfaces, units for forming such surfaces provided with twisting devices and control units for orientation of these surfaces. Orientation control units are made on base of gimbal mounts brought-out beyond craft hull. Each twisting device is made in form of hoop mounted on outer frame of gimbal mount for free rotation; it is engageable with electric motor. Units for forming reflecting surfaces are made in form of pneumatic systems with concentric pneumatic chambers and radial struts. Said struts are provided with flexible tubes with valves mounted at equal distances. Valves have holes. Built on said tubes are pneumatic cells in form of torus or spheres. Each pneumatic system is mounted on respective hoop and is communicated with compressed gas source through concentric hermetic groove found in hoop and in outer frame of gimbal mount.
EFFECT: reduced responsiveness of solar sail control; possibility of deploying solar sail according to preset program; reduction of mass per unit of surface.
4 cl, 11 dwg
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.