Carrier rocket sea surfacing and appropriate systems and methods
SUBSTANCE: invention relates to aerospace engineering and may be used in aerospace clustered rocket bodies. This system comprises space rocket with bilateral control surface to turn and to receive data on rocket section position on water surface and to control flight path. It comprises also launching site, rocket or rocket part launcher for launching for the first and second times. Means for vertical landing on water surface structure. Means for launching and means for varying the rocket orientation with tail forward before landing and reentry to the Earth atmosphere. Besides, it comprises means to rocket engine cutoff and means for primary start and restart of one or more rocker engines. Space rocket is launched with payload from the Earth, Said one or more engines are shutdown at accelerator stage. Upper stage is separated from accelerator stage at preset altitude. Accelerator stage orientation is changed to place movable landing platform on water surface. Data on landing platform is received to control accelerator stage path for displacement to landing platform. One or more rocket engines are restarted at accelerator stage before landing. Pocket part is vertically landed on said platform to carry said part on landing platform or to transit vessel.
EFFECT: vertical landing of rocket shuttle part on water surface landing platform.
20 cl, 2 dwg
The technical field to which the invention relates
The invention, in General, relates to a space launch vehicles, and more particularly to systems and methods for landing rockets on the sea and/or for the recovery of such launch vehicles on the route from the landing site.
The level of technology
Boosters with a rocket engine was used for many years for delivery into space man, as well as useful goods without people. The rockets brought the first men on the moon and launched into Earth orbit, many satellites, unmanned space probes, as well as deliver supplies and personnel to the Orbital international space station.
Despite the rapid development of manned and unmanned space flights shipping astronauts, satellites and other payloads into space continues to be expensive. One reason for this is that the most common boosters are used only once, and therefore they are called "disposable rockets-carriers" or "ELV". Advantages of launch vehicles reusable (RLV) are the possibility of entering space with little expense.
While the space Shuttle NASA is largely re-used, the recovery in shodnenskaya reusable components is costly and time consuming process which requires an extensive ground infrastructure. Furthermore, additional system space Shuttle, required for re-entry and landing, reduce the possibility of the payload of the space Shuttle. With increasing commercial pressure, there remains a need in the access to outer space with low costs for both manned and unmanned payloads.
Brief description of the drawings
Fig.1 shows a schematic diagram illustrating a General view of the flight of the space launch vehicle, which performs a landing on an offshore platform in accordance with an embodiment of the invention.
Fig.2 shows a block diagram illustrating a launcher of the rocket with the ground or other launch pads and landing of a space rocket on a sea platform in accordance with a variant implementation of the disclosure.
Detailed description of the invention
Certain aspects of the present disclosure is directed, in General, on a vertical landing with the engine running boosters reuse on the marine platform and corresponding systems and methods. Other aspects of the disclosure relate to the recovery of boosters reuse on the route from the place of landing of the sea or other place of landing. Certain items presented in the trail�following description and is shown in Fig.1 and 2 to ensure a complete understanding of various embodiments of the disclosure. For specialists in the art it will be understood, however, that other variants of implementation, having other configurations, layouts, and/or components, can be implemented in practice without several of the details described below. In particular, other embodiments of the disclosure may include additional elements, or they can lack one or more of the elements or properties described below with reference to Fig.1 and 2. In addition, several details describing structures and processes that are well known and often associated with space launch, and the launch and landing of space launch vehicles that are not represented in the following description to avoid unnecessary complications of the various embodiments of the disclosure.
In the figures the same numbers of reference positions denoted by identical or, at least, shared similar elements. In order to facilitate the description of any particular item, the highest significant digit or digits in any number of reference items refer to the drawing on which this element is first introduced. For example, element 110 is first introduced and described with reference to Fig.1.
Space launch vehicles usually run from pads located on the shore, through the corridors of the flight on which they podnimayts� over the ocean, over which passes a large part of their trajectories. This trajectory eliminates the potential risk to the public associated with the flight of the rocket, and leads to the falling stages of the accelerator in water. Landing on the water, however, makes reuse speed accelerator costly and difficult for a number of reasons. For example, sea water can be very corrosive to the components of the rocket. In addition, many components of rocket become very hot during use, and putting out these hot components cold sea water can lead to cracking and other forms of damage. The recovery and reuse of solid rocket stages after a water landing with a parachute feasible in practice, because solid-fuel rocket engine is nothing more than an empty body after combustion. Stage of the rocket with liquid fuel, however, are significantly more complex. As a result, few, if any, rocket stages with a liquid fuel re-use after a water landing.
Are the concepts of planting speed accelerator to the ground. These concepts include horizontal landing stages of the accelerator, as the plane, or vertical, using its own engine braking stage or on steam�Utah, or with another tool. All of these approaches, however, have limited flexibility because they require a landing site on earth for each azimuth start and potentially narrow the scope of the landing.
Other concepts have been proposed in which the degree of the accelerator is re-launching its rocket engines after separation from the upper stage (steps), and then flies back to the launch pad. Having arrived to the launch site, the degree of the accelerator or performs a horizontal landing on a runway, or vertical landing, using force or other means, such as a parachute. Both of these approaches, however, reduce the ability to recover the payload into orbit, since they require that the rocket was carrying a substantial load of rocket fuel to perform the maneuver with the flight back.
Fig.1 shows a schematic diagram illustrating the profile of the flight boosters reuse, which performs a vertical landing using its own engines on offshore platform in accordance with a variant implementation of the disclosure. In the illustrated embodiment, the implementation of multi-stage orbital launch vehicle 100 includes a first stage or step 110, the accelerator and a second or upper stage 130. Stage accelerator 110 may include � the structure of the intermediate compartment between the steps, includes an expandable aerodynamic surface 120 located towards the front end 114, and one or more rocket engines 116, located towards the tail end 112. Rocket engines 116 can include, for example, rocket engines with liquid fuel, such as engines with liquid oxygen/hydrogen, liquid oxygen/kerosene or motor RP 1, etc. In other embodiments, rocket engines 116 can include solid rocket fuel. As described in more detail below, the tail portion 112 of the stage 110 of the accelerator may also include a plurality of movable surfaces 118 governance (identified individually as surfaces 118a, 118b, etc. control), designed to manage both on the trajectory with increasing and trajectory with decreasing elevation stage 110 of the accelerator.
Although the upper stage 130 is installed over the stage accelerator 110 in the illustrated embodiment, in other embodiments, the launch vehicle 100 and its variants can have other configurations without going beyond the essence or scope of the present disclosure. For example, in one embodiment of the top step 130 and step 110 of the accelerator can be installed next to each other and attached to each other during the ascent with the help of COO�respective division. In another embodiment, the implementation of two or more stages 110 of the accelerator or their variants can be located around the top step 130 in the configuration of the type "link". In accordance with the present disclosure is not limited to a particular configuration of the booster shown in Fig.1.
In the illustrated embodiment, the implementation of the booster 100 run on the banks or other launch pads 140, and then she turns and flies over the ocean 102. In one aspect of this embodiment of the offshore platform 150 may include a station 152 broadcasts, intended for the transmission of its position in the booster 100 in real time. This information allows the booster 100 and/or 110 degrees of the accelerator constantly check and/or adjust your flight path so that it was focused on the platform 150. If the platform 150 is a freely drifting vessel, the platform 150 may also include the prediction block in the platform (for example, an appropriate processing device, storage device and corresponding computer executable instructions) that automatically predicts the future position of the platform 150 on the basis of various existing conditions, such as the strength and direction of ocean currents, force and wind direction, current speed and direction of drift, etc. for Example, the block prediction of the position of the platform may be configured to predict the position of the platform at the expected time of landing of the rocket. In addition, the station 152 broadcast can pass this information on booster 100 and/or a step 110 of the accelerator in real time so that the launch vehicle 100 and/or the degree of accelerator 110 can use this information to regulate their flight paths and better targeting the location of the landing. After cutoff of the engine accelerator (OUESSO) at high altitude stage accelerator 110 is separated from the upper stage 130 and continues flying along a ballistic trajectory. The engine or engines of the upper stage 132 (e.g., a liquid-propellant engines) can then be launched and can provide flight upper stage 130 at a higher trajectory 134 orbit or to other destinations. As step 110 of the accelerator re-enters the earth's atmosphere, it changes the orientation so that the tail end 112 is installed in the direction of travel, and plans in the direction of the marine landing platform 150. In another embodiment, the implementation stage accelerator 110 may re-enter the atmosphere nose forward and then edit�th orientation on the orientation of the tail forward just before landing. In still another embodiment, one implementation of the landing of the rocket engine and/or corresponding structures landing gear can be installed on the front end 114 of the stage 110 of the accelerator so that the notch 110 of the accelerator could re-enter the atmosphere nose forward and then perform a landing in a nose down orientation.
Depending on the specific trajectory of the launch of the offshore platform 150 may be located at a distance of hundreds or more miles from located on the shores of the pad 140. As the degree of accelerator 110 is lowered in the direction of the offshore platform 150, the degree of accelerator 110 can regulate the course of planning, aiming at the platform 150, based on the data in the platform, taken from station 152 broadcast. In addition, or alternatively, offshore platform 150 may include submerged or partially submerged drive system (having, for example, propellers or other drive device) for maintaining the platform 150 in position or displacement of the platform 150 in accordance with the need to control the drift and/or deflect the accelerator. One or more boats with ropes can also be used to hold the platform 150 in the desired position or displacement of the platform 150 in with�testii with the need to control the drift and/or changes the trajectory of the accelerator.
As step 110 of the accelerator decreases towards offshore platform 150, the degree of accelerator 110 may control the course of their planning, using surface 118 aerodynamic controls located on the tail end 112, and/or the expandable surface 120 controls located towards the front end 114. In one aspect of this embodiment of the expandable surface 114 of the management can include aerodynamic surfaces that open up or unfold in an outward direction, in the form of, for example, shuttlecock, to create aerodynamic thrust behind the center of gravity (CG) stage 110 of the accelerator, which helps to stabilize the tread 110 of the accelerator in the orientation of the tail forward. In another aspect of this variant implementation is made with the possibility of movement of the surface 118 of the aerodynamic controls located towards the tail end 112 of the accelerator 110, may include a bidirectional control surface, which can control the height and/or path stage 110 of the accelerator as during the ascent, when the launch vehicle 100 moves in the forward direction, and decrease when the degree of accelerator 110 moves in the direction of the tail forward to sailing on the sea platform 150. Accordingly in one aspect Danno�of embodiment of the surface 118 of the aerodynamic control are bi-directional, supersonic control surfaces. In other additional embodiments, the corresponding system of the parachute can be deployed from, for example, the front end 114 of the stage 110 of the accelerator to reduce and/or otherwise control the rate of descent during all or part of the reduction.
After the step 110 of the accelerator will go down in the appropriate position above the platform 150 (e.g., in some embodiments from approximately 100,000 feet to about 1000 feet, or in other embodiments from about 10,000 feet to about 3000 feet), it re-starts the engine 116 of the accelerator to slow the speed of his decline. Step 110 of the accelerator then performs a vertical landing with the engine running on the platform 150 low speed. For example, step 110 of the accelerator may slow down the speed reduction of approximately 60 feet per second to approximately 1 foot per second or less and can be planted on a landing platform 150 using cardan suspension engine 116 of the accelerator and/or the engine management system height to control the height and/or position of the stage 110 of the accelerator during landing. In one embodiment, the implementation stage 110 of the accelerator can be planted on appropriate shock absorbing chassis. In other vari�nth implement other means of planting can be used for the landing stage 110 of the accelerator on the marine platform 150 in accordance with the present disclosure.
In another embodiment, the implementation of one or more jet engines (not shown) can be properly installed on the tail end 112 or in another area of the stage 110 of the accelerator to perform all or part of the maneuvers in a vertical landing. Jet engines can be started during the decline stage of the accelerator and can be used in combination with or instead of repeatedly running engine accelerator 116. Jet engines can be more efficient in fuel consumption than the engines 116 of the accelerator and, as a result, can provide more flight time and better control to step 110 of the accelerator during landing on the platform 150. In one embodiment of the jet engines may be used in combination with an appropriate system of parachute, which expands and decelerates the step 110 of the accelerator before starting jet engines.
In one embodiment, the implementation of the offshore platform 150 may be a free-floating, designed to voyage across the ocean barge with appropriate deck intended for the landing and transportation of stage 110 of the accelerator. In other embodiments, the platform 150 may be part of a more complex vessel, such as a semisubmerged platform with underwater Digi�ate, to minimize or at least reduce movement of the deck and hold it in a fixed or relatively fixed position. In the variant of implementation, based on barge offshore platform 150 can be towed back to the pad 140 that is located on the beach, or other port after landing for repair and/or recovery for reuse. In one embodiment, the implementation of the offshore platform 150 can be towed with a tug boat or other suitable vessel. In other embodiments, offshore platform 150 may include its own drive system in motion for the transportation stage 110 of the accelerator back on the pad 140 or in another port.
There are many benefits associated with the options of implementing the present disclosure described above with reference to Fig.1. For example, the reconstruction stage 110 of the accelerator after landing on offshore platform reduces the costs associated with the launch of a multistage launch vehicles. In addition, by performing a vertical landing using its own engines, the degree of the accelerator is restored in such a way that minimizes or at least reduces the amount of restoration work required to re-use�tion. In addition, embodiments of the disclosure described above, can improve the functional flexibility of launch vehicles, as live ocean platform 150 may move to another area of the ocean when changing the azimuth of the launch of mission and/or upon change of the place of intended landing of the flight. In addition, live on the ocean platform 150 may even be moved to other parts of the world to support launches from other sites (e.g. other pads located on the Bank). In addition, to start with pads located on the shore, the launch vehicle 100 can also be run from the sea to live on the ocean platform or vessel and then can land in range of the landing live on the ocean platform 150. Such options for implementation may be preferred for Equatorial launches from offshore platforms to increase payload. Alternatively, in other embodiments, the booster 100 can be started with moving on the ocean platform, and then the accelerator 110 can be restored as a result of performing a vertical landing using its own engines on land.
Embodiments of the disclosure described above, can also increase opportunities for the delivery of the payload of the launch vehicle 10, enabling more efficient flight level 110 of the accelerator, or at least using a very efficient path, when she re-enters the atmosphere and is moving in the direction of the platform 150. Capacity useful load increase, because it is not required to contain the propellant in stage 110 of the accelerator for the flight back to the landing site, located on land. In addition, offshore platform 150 may be installed in any location, which will identify the landing stage 110 of the accelerator after separation from the upper stage 130. The variants of the implementation disclosed here, can also reduce or eliminate public concerns about safety associated with the reversal of the trajectory of the stage 110 of the accelerator for landing on land.
Embodiments of the above disclosure also allow to solve the problem of transportation of a stage 110 of the accelerator back either to the pad 140, located on the shore, or to another facility situated on land recovery. More specifically, the degrees of the accelerator boosters are usually very large in size, and as a result of their transportation fully assembled can be a significant problem in relation to logistics and �of atrat. If the degree of the accelerator will be planted on the land, the problem of transportation speed accelerator back or on the pad, or to the place of recovery must be resolved, and transportation by land objects with dimensions speed accelerator is a problem from the point of view of logistics and Finance. In contrast, transport in the ocean is a cost-effective means of transporting large loads, such as degree of the accelerator, over long distances. Offshore platform 150 in accordance with the present disclosure may be towed back to the Harbor next to the pad and unloaded to restore and re-use at relatively little cost.
Although in Fig.1 shows a variant implementation of the disclosure in the context of the recovery stage of the accelerator, the present disclosure can also be used to recover a returned funds from orbit with accuracy, provide a vertical landing using its own engines. One advantage of this approach is that it could provide the possibility of installing an offshore platform 150 in any area of the ocean or other body of water (e.g., Strait, lake, etc.) suitable for planting by the returning delivery vehicles. In addition, many �viewwise by the sea of platforms can be deployed worldwide in predetermined locations to provide landing zones for emergency situations if you want to aborted the mission.
Fig.2 illustrates a procedure 200 in the form of a block diagram the sequence of operations of way of launching and landing of a space booster, such as booster in accordance with a variant implementation of the present disclosure. In one aspect of this embodiment, the procedure 200 can be embodied for booster 100, described above with reference to Fig.1. In other embodiments, procedure 200 or part thereof may be used in other types of launch vehicles including launch vehicles, pornitalia launch vehicles, launch vehicles into space and interplanetary funds, etc.
In block 202, the procedure begins with the engine start-up accelerator and lift with pads (e.g., pads on land, such as a pad located on the Bank). As described above, in other embodiments, the mission can begin with lifting pads located on the sea, such as a floating platform, barge, ship or other vessel. In block 204 occurs, the shutoff of the engine accelerator at a given height. In block 206, the upper stage separates from the degrees of the accelerator and the engine or engines of the upper stages are launched.
In block 208 the degree of the accelerator, change the orientation, following its ballistic t�actorii after separation from the upper stage. More specifically, the degree of the accelerator, change the orientation so that it moves with the tail forward. In one embodiment, the implementation of the change in the orientation degree of the accelerator can be performed with the use of deployable aerodynamic surfaces (e.g., expanding surfaces), which continues outward from the front end stage of the accelerator for receiving the thrust force behind the CG-stage accelerator. In other embodiments, the thrusters (e.g., rocket propulsion, such as propulsion to the hydrazine can be used in addition to or instead of the aerodynamic control surfaces to change the orientation degree of the accelerator. For example, if the change in the orientation degree of the accelerator takes place in outer space, where the aerodynamic control surfaces are ineffective, then the thrusters can be used to change the orientation degree of the accelerator.
In block 210 the surface of the aerodynamic thrust and/or control unfold before or during re-entry means of delivery in the Earth's atmosphere. In block 212 the degree of accelerator re-enters the atmosphere and establishes contact with moving by sea landing platform. Alternatively, the booster can make contact with moving by sea landing platform before �simple entrance or can be in constant contact with moving by sea platform during the whole flight. In block 214, the degree of accelerator plans or, in other words, it follows a ballistic trajectory towards moving by sea landing platform.
In block 216 the decision procedure determines whether to adjust the degree of the accelerator to properly install speed accelerator on moving by sea platform. If not, the procedure goes to block 220 and the degree of accelerator continues to plan towards moving by sea platform. If you need to adjust the trajectory planning, the procedure goes to block 218 and actuates the aerodynamic control surface to change the trajectory planning stage of the accelerator. Alternatively, or in addition to changing the trajectory planning stage of the accelerator, the procedure can also adjust the position of the landing platform using, for example, the drive system in the movement associated with boarding platforms or by towing of the platform.
After adjusting the trajectory planning and/or position of the landing platform, the procedure goes to block 222 of the decision to determine whether the situation degree of the accelerator is installed on a landing platform, suitable for the preparation to the final stage of landing. If not, the procedure returns to block 216 acting� decisions and repeats. Once the tool is installed in the appropriate position over a landing platform for the preparation to the final landing procedures, the procedure goes to block 224 and re-start the engines of the accelerator. In block 226, the booster performs a vertical landing using its own engines on the marine platform, and the flight phase of the procedure ends.
However, in one embodiment, the implementation of the procedure 200 may continue at block 228 to move the platform and steps of the accelerator back to the pad or to another port for recovery and reuse. In block 230, the degree of the accelerator is reduced in accordance with the need and establish the new booster. From block 230, the procedure returns to block 202 and is repeated for the new rocket.
In a particular embodiment of the offshore platform can be positioned so that it improves and/or optimizes the separation of the second stage booster, for example, as in azimuth, and the distance from the pad. For example, in at least some cases, the possibility of moving moving by sea platform allows you to extend the range of available locations where pad accelerator is separated from the rest of the booster, the pic�ol'ku landing site accelerator is not very strictly limited. The ability to control the trajectory by reducing the accelerator can further extend the range of available landing areas.
In any of the described variants of implementation after landing rocket total process may include additional steps that contribute to the rapid return of the rocket to work. For example, the booster can be transferred relatively slowly moving by sea platform on a faster surface vessel to reduce its translation back onto the pad. In addition to or instead of transferring the use of the rocket can be recovered during its transportation from the place of landing to the pads. Aspects of both of these properties is described further below in the context of the recovery of the sea launch vehicle. In other embodiments, the implementation of specific aspects of these properties (for example, recovery of the launch vehicle and the route from the place of landing) can be applied to other configurations of the restoration, including the restoration of land.
In a particular embodiment of the booster (e.g., a reusable system accelerator first stage or RBS) direct and/or offline transferred to a safe state after the marine landing on a landing platform and before the treatment & �is designed to remedy. Autonomous action security can include ventilation of tanks for rocket fuel and cylinders displacing gas, and clean any aerodynamic surface. The tool can then be moved to a separate, smaller ship for a more rapid return to the pad on the shore or to the site overloading. In another embodiment, the implementation of the tool may be secured to the deck of the landing platform and the platform can be towed or can be moved using its own drive back to the pad on the shore or to the site overloading. In each case, the tool can be moved using a marine crane (or other suitable device) for fixing tools output in vertical or in horizontal position for transport over the ocean, and may be unloaded onto a truck in the dock to return to the enterprise for processing the output on the pad.
During the movement, as well as the enterprise processing means output booster can be processed for the next run. Activities associated with reuse, which is usually performed before each launch, you can include the items of maintenance (if any), cleaning, recharging of cylinders of vitess�appropriate gas, recharge electric batteries, recovery of materials thermal protection system, if required, and/or functional testing of pneumatic, hydraulic subsystems, and avionics subsystems. While on the route or in the enterprise for processing output means, output means can be docked with nonrecoverable upper stage, which can be pre-assembled with a payload and the payload fairing. In other cases, the booster can be directly connected to the payload module. Through periodic intervals can also be performed basic maintenance, such as engine overhaul.
During the above operation, when processing, if the system includes one offshore platform and it is used for transportation means output back to the pad on the shore, the platform can then be re-installed in the landing zone after unloading means output and will be ready for planting second output means, while the first tool is on the way to the pad. If the whole system includes two offshore platforms, one offshore platform may remain in the landing zone between flights, while the other is a Ter�Xia Bank. In still another embodiment, one implementation of the system may include two booster, one marine landing platform and a separate vessel that transports the tool from the platform to the pad, which also allows one the landing platform to stay in the landing zone between flights. Individual vessel may include a ship or aircraft in certain embodiments of the implementation.
In any of the previous embodiments may be made of any suitable aspect of the recovery process during transportation funds withdrawal, provided that, for example, such a process can be successfully completed in marine conditions and accordingly will be docked with the subsequent processes.
Based on the above it should be understood that the specific embodiments of the invention have been described herein for illustration, but that various modifications may be made without deviation from the essence and scope of the various embodiments of the invention. For example, although various embodiments of the present disclosure have been described above in the context of landing booster on a sea, in other embodiments, the system and method described herein may be used for landing booster on other water wantin�spaces, including, for example, lake, Bay, ocean, Strait, or perhaps even a great river. In addition, while various advantages associated with certain embodiments of the implementation of the disclosure have been described above in the context of those embodiments, other embodiments of can also exhibit such advantages, and not all variants of implementation must exhibit such advantages to fall within the scope of the invention. In accordance with this invention is not limited by anything except the appended claims.
1. A method of controlling a space launch vehicle, including payload, is placed on the top step mounted on the stage of the accelerator, characterized by the fact that:
start a space rocket from the Earth, and the launch of a space rocket from the earth involves running one or more rocket engines in the degree of accelerator to launch a space rocket from the launch pad on Earth with the orientation of the nose forward
disable the specified one or more rocket engines at the degree of the accelerator;
separating the upper stage from the stage of the accelerator at a given height;
change the orientation degree of the accelerator in a tail first orientation;
place mobile landing platform �and water surfaces,
take the information about the position of the landing platform and control the trajectory speed accelerator to move to the landing platform when the orientation of the tail forward on the basis of said position information;
perform a restart of the specified one or more rocket engines on the steps of the accelerator before planting, and
perform landing the space launch vehicle on a specified landing platform on the water surface, while landing the space launch vehicle includes a vertical landing stage of the accelerator on the platform with the orientation of the tail forward, providing the thrust from re-running one or more rocket engines.
2. A method according to claim 1, wherein landing the space launch vehicle includes a vertical landing speed accelerator on a floating platform on the water surface.
3. A method according to claim 1, further characterized by the fact that many times using at least a part of a space launch vehicle.
4. A method according to claim 1, wherein additionally:
transported by space rocket on a specified mobile landing platform to the object recovery;
restore at least part of the space launch vehicle on the subject of recovery; and
use at least a portion of the rocket-wear�I after recovery.
5. A method according to claim 1, which further overload the reusable part of the space launch vehicle from the landing platform to transit the ship, landing platform left on the water surface for receiving subsequently launched booster.
6. A method of transporting a payload into space, characterized by the fact that:
connect the payload stage rocket accelerator, wherein the step of the accelerator contains a plot of the front end, located at a distance from the area of the caudal end;
place a floating platform in a certain position on the water surface;
start one or more rocket engines mounted in the direction of the section of the caudal end-stage accelerator, and launch a rocket into outer space in the orientation of the nose forward;
disable the specified one or more rocket engines,
separate the payload from degree of the accelerator;
after separation change orientation degree of the accelerator with the orientation of the nose forward of the tail first orientation;
accept the position information of the floating platform and control the trajectory speed accelerator to move to a floating platform at a tail first orientation based on the specified position information;
breach�t re-start the specified one or more rocket engines before landing, and
perform the landing stage of the accelerator on a floating platform in the orientation of the tail forward, landing stage accelerator includes a vertical landing stage of the accelerator on the platform with the orientation of the tail forward, providing the thrust from re-running one or more rocket engines.
7. A method according to claim 6, in which
disable the specified one or more rocket engines includes eliminating one or more rocket engines before the orientation change degree of the accelerator with the orientation of the nose forward of the tail first orientation.
8. A method according to claim 6, which additionally:
after turning off the specified one or more rocket engines and follow a ballistic trajectory deploying aerodynamic surface control of degree of the accelerator to facilitate the change in the orientation degree of the accelerator with the orientation of the nose forward of the tail first orientation.
9. A method according to claim 6, which additionally:
after turning off the specified one or more rocket engines and follow a ballistic trajectory deploy one or more extending control surfaces from the area of the front end stage of the accelerator to facilitate the change in the orientation degree of the accelerator from the orientation of the nose forward on Oriental�Yu caudad.
10. A method according to claim 6, which additionally:
perform operations with one or more engines, reaction control system, installed at the degree of an accelerator to facilitate the change in the orientation degree of the accelerator with the orientation of the nose forward of the tail first orientation.
11. A method according to claim 6, which additionally:
move the aerodynamic surface control in the degree of the accelerator so as to at least partially control the trajectory of the degrees of the accelerator to a target platform based on the information about the position of the platform, adopted from the platform;
move the aerodynamic surface control in the degree of the accelerator so as to at least partially change the orientation degree of the accelerator from the orientation of the nose forward of the tail first orientation.
12. System for providing access into the space that contains:
means for launching the launch vehicle from the launch site for the first time;
tool for planting at least part of the booster on the structure on the water surface; and
means for running at least part of the booster with the specified launch pad a second time.
13. A system according to claim 12 in which the means for planting incl�et a means for vertical landing, at least part of a space rocket on a floating platform.
14. A system according to claim 12 in which the means for launching includes means for launching the rocket in the orientation of the nose forward, the system further comprises means for changing the orientation of the launch vehicle with the orientation of the nose forward of the tail first orientation before boarding, and the means for landing includes a means for landing in a tail first orientation.
15. A system according to claim 14, in which the space launch vehicle includes one or more rocket engines, wherein the means for launching includes means for including rocket engines and launch of the carrier rocket in the orientation of the nose forward and the system further comprises:
a means to disable rocket engines;
the means for changing the orientation of the specified means to start with the orientation of the nose forward on the orientation of the tail first before planting; and
means for restarting one or more rocket engines, when this booster is in the orientation of the tail forward to slow down the booster, wherein the means for landing includes a means for landing with a tail first orientation, and with respect to the specified one or Bo�more rocket engines.
16. A system according to claim 12 in which the means for launching includes means for launching a rocket into outer space in the orientation of the nose forward, wherein said system further comprises means for changing the orientation of the launch vehicle with the orientation of the nose forward of the tail first orientation, wherein the orientation change to re-enter the Earth's atmosphere.
17. A system according to claim 12, in which the space launch vehicle includes one or more rocket engines, wherein the means for launching the launch vehicle includes a means to include rocket engines and launch of the carrier rocket in the orientation of the nose forward, wherein said system further comprises:
a means to disable rocket engines;
the means for changing the orientation of the specified means to start with the orientation of the nose forward on the orientation of the tail forward, and
means for controlling the flight path of the rocket as it moves to the specified structure on the surface of the water.
18. A system according to claim 12, in which space the booster is arranged to receive position information of the specified structures on the surface of the water and adjust the trajectory of the space launch vehicle based, at least in part, the specified INF�rmacie.
19. A system according to claim 12, in which the space launch vehicle includes a bidirectional control surface.
20. A system according to claim 12, in which the space launch vehicle includes an expandable control surface, space booster is made with the ability to deploy the control surface after launch and before landing.
SUBSTANCE: invention relates to space and missile equipment and can be used for final carrier rocket stages. A space and missile system (SMS) includes a carrier rocket with a final stage with an outer housing compartment with an intermediate power support frame with outer and inner frames connected to each other by means of fastening elements; a space vehicle with the main fairing with an end frame. The outer diameter of the intermediate support power frame corresponds to the diameter of the end frame of the main fairing.
EFFECT: invention allows attachment of different standard sizes of main fairings with carrier rockets without increasing the time for assembly and preparation for SMS launching.
SUBSTANCE: invention relates to systems for payload delivery to upper air and higher. Proposed system (1) comprises tubular rocket launching cart (2) with friction drives of cable-rope path (26) displacing below two-pin hinge (63) secured to ground and lifted to coaxial portable tube (124, 143) extending to three main fastened cables/ropes (27), their weight being equalized by balloons (164). Then, said cart displaces to connector station (166) retained above the ground in stratosphere by two secondary cables/rope (184) suspended from fastening frame (162) to tensioning of said balloons. Said cart is retained by end grip (196) of said cart directed in two secondary and two tertiary cables/rope (186) to be lifted by bottom lifter (198) guides by secondary cables. Said bottom lifter is retained by top lifter (168) suspended from fastening frame of tensioning balloons. Said cart hooked by lifting ring (183) guided by secondary cables/ropes upward revolves in necessary direction with release of rocket and, fact, recoilless ejection during free fall of said cart downward with engine ignition at safe distance.
EFFECT: safe, non-polluting reusable system.
50 cl, 67 dwg
FIELD: aircraft engineering.
SUBSTANCE: rocket cryogenic upper stage (RCUS) designed according to the tandem layout comprises a fuel tank with an instrument compartment and transitional system for fastening of a spacecraft, an oxidizer tank (OT), intertank spacer, RCUS mid-flight engine (MFE), an intermediate compartment, fire and explosion prevention system, thermal mode maintaining system with the unit of demountable connections of communication with the land equipment and separable inlet pipelines, manifolds for purging of stagnant zones and device for maintaining of the thermal mode of the zone and RCUS equipment, a sealing diaphragm, the detachable head fairing (HF) with windows for detachment of the fire and explosion prevention system and devices for maintaining of the thermal mode of gases for purging of RCUS zone, additional thermal insulation of RCUS zone, a part of separable inlet pipes of manifolds with demountable joints and the unit of demountable connections for communication with the land equipment, an intertank spacer, conjugated with the intertank frame for fastening OT with MFE and conjugated to the top spacer of the separated intermediate compartment with the units of connection and separation with US and HF.
EFFECT: invention allows to improve fire and explosion safety of upper stage.
FIELD: aircraft engineering.
SUBSTANCE: invention can be used for control of travel of space liquid-fuelled rocket (SLFR). After the command for disabling of the mid-flight engine (MFE) of the burnt-out stage MD is switched to the reduced-thrust stage and MD is completely disabled, the rolling travel of the rocket is controlled using two pairs of gas nozzles, the moment of MD final disabling is forecasted, one of pairs of gas nozzles is enabled before the forecasted moment of MD final disabling for creation of the control rolling moment, the pair of gas nozzles are disabled at forecasted moment, and the value of the period of work of pair of gas nozzles is determined before the flight depending on the moment of inertia of the rotating part of the turbo-pump unit with the allowance for attached mass of fuel components with reference to the axis of rotation, an absolute value of the roll moment, created by each pair of the gas nozzles at their enabling, an absolute value of angular speed of rotation of the turbo-pump unit rotor at the reduced-thrust stage, angle between the axis of rotation of the turbo-pump unit rotor and the longitudinal axis of the rocket.
EFFECT: invention allows to improve safety of SLFR flight.
FIELD: physics, navigation.
SUBSTANCE: group of inventions relates to interorbital, including interplanetary, flights of rocket propelled spacecraft. A method of constructing an optimal spacecraft trajectory is based on solving a two-point boundary value problem of the Pontryagin maximum principle and taking into account characteristics of the macro- and microstructure of the cost function. The latter can be the time of flight or fuel consumption during flight. Analytical bases for efficient search of initial domains of values of Lagrange multipliers at each iteration are established. This facilitates the construction of a series of sub-optimal solutions which converge to an optimal solution. A corresponding algorithm yields the optimal solution last or, in case of unattainability thereof (due available resources of the spacecraft) a solution close to optimal. An electronic processor for implementing the method and a spacecraft with said processor are also disclosed.
EFFECT: faster operation, improved convergence, low qualification requirements and wider field of use of the disclosed algorithm and accompanying equipment.
16 cl, 7 dwg
FIELD: physics, atomic power.
SUBSTANCE: invention relates to atomic power engineering and space-rocket engineering. The spacecraft nuclear propulsion system comprises a heater - gas-cooled nuclear reactor, a cooler, a recuperative heat exchanger, a pipe system with a gaseous working medium, coaxial turbine-compressor-electric power generator, electric jet engines, an automatic control system with measurement and control means. The number of loops of the turbine-compressor-electric power generator with equal electric power is a multiple of two with opposite direction of rotation of rotors of the turbine-compressor-electric power generator in each pair, wherein the pipe system connects the output of the heater - gas-cooled nuclear reactor with the input of each turbine, and the output of the turbine with the input of the channel of the heated gaseous working medium of its recuperative heat exchanger, the output of the channel of the heated gaseous working medium of the recuperative heat exchanger with the input of its cooler, the output of the cooler with the input of its compressor, the output of the compressor with the input of the channel of the cold gaseous working medium of its recuperative heat exchanger, the output of the channel of the cold gaseous working medium of each recuperative heat exchanger with the input of the heater - gas-cooled nuclear reactor.
EFFECT: high efficiency and reliability of the spacecraft nuclear propulsion system.
19 cl, 2 dwg
SUBSTANCE: invention relates to space engineering and can be used in carrier rockets.. Proposed rocket comprises head unit with payload, parallel separable rocket stages with multichamber engines with fuel tanks shaped to torus, tapered tail, short central body at first stage, single trough-like nozzle at second stage, bottom part composed of outer and inner cones composed by outer and inner surfaces of short central body shell and inner surface of single trough-like nozzle shell. Fuel tanks and single trough-like nozzle are arranged inside short central body between first-stage tanks.
EFFECT: decreased bottom resistance, higher specific pulse.
5 cl, 9 dwg
SUBSTANCE: invention relates to space engineering, particularly, to astronaut operation in weightlessness. Proposed holder comprises retainer composed by wire (made of afterflow material) in non-metallic sheath, ring at retainer end in diameter comparable with sized of fingers of inflated space-suit glove, lever with opening in diameter comparable with retainer diameter.
EFFECT: higher safety of articles retention in open space.
SUBSTANCE: invention relates to space engineering, particularly, to astronaut operation in weightlessness. Proposed holder comprises retainer composed by wire (made of afterflow material) in non-metallic sheath, rings at retainer end in diameter comparable with sized of fingers of inflated space-suit glove.
EFFECT: higher safety of articles retention in open space.
FIELD: engines and pumps.
SUBSTANCE: pulse is obtained by ejection of gasified liquid residues of unused components of rocket propellants (RP). Pulse is generated by combustion of unused components of rocket propellants (RP) on rocket gas engine combustion chamber. Volume of unused propellant residues is limited to divide a second heat carrier mass flow rate into parts, one being fed in tank section confined by the screen while another portion being fed into tank second part. Amount of fed heat carrier is defined proceeding from evaporation of residual propellant component drops. Device for withdrawal of separable carrier rocket section comprises oxidiser and propellant tanks, tank supercharging system, rocket gas engine with feed and gasification systems. It incorporates feed lines with acoustic radiators (calculated proceeding from minimum mass loses for gasification by preset amounts of propellant and pressure). Said separation screen is calculated proceeding from surface tension force.
EFFECT: reduced power consumption for gasification.
3 cl, 4 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.
FIELD: rocketry and space engineering; designing artificial satellites.
SUBSTANCE: proposed spacecraft has modules where service equipment is arranged and modules where target equipment and command and measuring devices are located. Optical devices of target equipment of infra-red range with cooled elements are mounted in central module. Radio equipment of on-board repeater is arranged in side modules whose position is changeable relative to position of central module. Optical and command and measuring devices are mounted on one frame at reduced coefficient of linear thermal expansion; they are combined with central module through three articulated supports. Cooled elements of optical devices are connected with radiators located beyond zone of thermal effect; service equipment module is provided with solar batteries having low dynamic effect on accuracy of spacecraft stabilization. Besides that, this module is provided with plasma engine whose working medium excludes contamination of said optical devices.
EFFECT: enhanced accuracy of spacecraft stabilization; electromagnetic compatibility of systems.
FIELD: rocketry and space engineering; adapters for group launch of spacecraft.
SUBSTANCE: proposed adapter has body consisting of two parts: one part is made in form of load-bearing body with platform for placing the spacecraft on one end and with attachment frame on other end; other part is made in form of load-bearing ring secured on payload frame and provided with attachment frame. Attachment frames of load-bearing body and load-bearing ring are interconnected by means of bolted joints fitted with two rubber washer shock absorbers each; one of them is mounted between surfaces of attachment frames to be coupled and other is mounted between opposite surface of attachment frame of load-bearing body and metal washer laid under bolt head. Diameter of metal washer exceeds diameter of rubber washer shock absorber; spacecraft attachment units are secured on platform of load-bearing body by means of bolted joints with rubber washer shock absorbers mounted between platform surfaces to be coupled and spacecraft attachment units.
EFFECT: reduction of dynamic vibration and impact loads due to extended range of varying dampening properties of adapter.
6 dwg, 1 tbl, 1 ex
FIELD: future space engineering; interstellar flights.
SUBSTANCE: proposed method is based on use of reactive thrust of spacecraft rocket engines in their maneuvering in gravity field of black hole. Kerr (rotating) black hole, i.e. its ergosphere may be selected for the purpose. Several separate spacecraft are directed in succession to gravity field of black hole ensuring stable exchange of information among them (for example, by radio or light channel). Provision is made for acceleration of spacecraft to relativistic speeds and obtaining information on effect of such speeds and accelerations on physical processes, equipment and living beings (at safe flying out of sphere of influence of black hole), as well as verification of theories of black holes.
EFFECT: enhanced efficiency.
FIELD: rocketry and space engineering; upper stages of launch vehicles injecting payloads from reference orbit into working orbits.
SUBSTANCE: proposed cryogenic stage includes cruise engine, oxidizer tank, toroidal fuel tank, inter-tank compartment, truss for connection with payload and truss for connection with launch vehicle. Toroidal fuel tank is made in form of lens in cross section with bottoms changing to frames. Tank is coupled with said trusses and inter-tank compartment through outer frame forming load-bearing system for taking-up external inertial loads.
EFFECT: reduction of total longitudinal clearance and mass of cryogenic stage; increased zone of payload under launch vehicle fairing.
FIELD: rocketry and space engineering; scientific and commercial fields.
SUBSTANCE: proposed method includes placing payloads on injection facility, launching the launch vehicle, separation of injection facility from launch vehicle and injection of injection facility into geocentric orbit where said payloads are separated from injection facility. Main payload is placed on injection facility directly of body of accompanying payload; this body combines its functions with functions of main load-bearing member of adapter system for placing the main payload. After separation of injection facility from launch vehicle, additional acceleration of injection facility is performed and injection facility is injected into reference orbit and then it is shifted to geocentric orbit where main and accompanying payloads are separated. Accompanying payload is separated from injection facility after main payload is at safe distance without waiting for complete turn of main payload. Spacecraft in facility injecting the artificial satellites into geocentric orbit are placed in succession on injection facility beginning with lower one. Main payload in form of one or several spacecraft is placed on body of lower spacecraft through separation device. Body of lower spacecraft combines its functions with functions of adapter load-bearing member for placing the main payload.
EFFECT: increased mass ratio of launch vehicle and injection facility; extended functional capabilities.
3 cl, 2 dwg
FIELD: space engineering; spacecraft for descent in atmosphere of planet.
SUBSTANCE: proposed spacecraft has case with foldable wings and/or stabilizers provided with deployment mechanisms. In folded state at deceleration of spacecraft in atmosphere, said wings and/or stabilizers are covered with separable frontal heat shield which is oval in shape in projection on plane perpendicular to longitudinal axis of spacecraft. Side surfaces of tail section of spacecraft case with wings and/or stabilizers (and some other members) may be covered with separable aerodynamic flaps which form conical surface. After deceleration at initial stage of descent, shield is separated and wings (stabilizers) deploy to working position. Proposed spacecraft has high aerodynamic properties and is provided with reliable protection against aerodynamic and thermal loads at deceleration at high supersonic flight speeds.
EFFECT: low cost of servicing.
4 cl, 13 dwg
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: 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
FIELD: space engineering.
SUBSTANCE: proposed method includes transportation of space launch vehicle to launching position, preparation for launch, raising the space launch vehicle to preset altitude by carrier-aircraft, separation from carrier-aircraft, stabilization of space launch vehicle and starting the engine plant of first boost stage. Space launch vehicle is transported to launching position in transportation-and-operation container. Then, container is transferred by means of crane to erection trolley, detachable compartments are dismantled and space launch vehicle is transported to carrier-aircraft. Space launch vehicle is secured to carrier-aircraft by means of locks of carrier-aircraft. Space launch vehicle is equipped with boost stages with solid-propellant engine plants, stabilization unit and units for attachment of launch vehicle to carrier-aircraft. It is also equipped with separable tail fairing and lattice stabilizers made in form of cylindrical panels which are secured on it. After bringing the space launch vehicle to preset altitude, locks of carrier-aircraft are opened by command and lattice stabilizers of tail fairing are opened simultaneously. After preset pause, before separation of space launch vehicle, tail fairing with lattice stabilizers is separated from space launch vehicle. Proposed method makes it possible to reduce launch mass and ensure stabilization on flight leg of safe distance from carrier-aircraft till moment of start of 1st stage engine plant.
EFFECT: extended field of application.
7 cl, 5 dwg