RU2479469C1 - Gliding spaceship (versions) with folding nose cone and method of controlling its descent to airfield - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to aerospace engineering. Proposed spaceship comprises heat-resistant airframe, rotary outer wings, aerodynamic and jet rudders, plug-in modules at rocket propellant tanks, and rocket carrier sustainer power plant module. In compliance with first version, spaceship comprises rocket carrier sustainer power plant module nozzle arranged at spaceship airframe front part and oriented forward relative to said airframe. In compliance with second version, spaceship comprises habitable cargo-and-passenger compartment with space pilot workstation at nose dome communicated with separable conformal orbital compartment via sealed hatch. Proposed method of control comprises inputting corrections to gliding spaceship orientation and its path by aerodynamic and gas-dynamic rudders in compliance with preset law of controlling descent and landing at airfield runway. Control effects using aerodynamic forces are effected in pitch channel be deflecting elevators.Those in bank channel are effected by asymmetric deflection of outer wings. Those in track channel are performed by combined deflection of elevator and asymmetric deflection of outer wings. Besides, additionally, synchronous outer wing elevation or sinking may be used in air speed channel.

EFFECT: decreased overall dimensions.

10 cl, 10 dwg

Description

Technical field

The invention relates to rocket and space technology, and more specifically to the areas of launch vehicles and space transport vehicles of repeated use, and is aimed at improving the layout of such vehicles, as well as the way to control the return to the airport of their reusable modules.

The level of technology.

The task of multiple use of a spacecraft with a marching propulsion system of the second stage of the launch vehicle and its instrumentation was first successfully solved in practice in the early 80s of the 20th century on the Space Shuttle Space Shuttle, created in the USA. The set of basic technical solutions that provide this technical result is described, for example, in patent US 3702688 from 11/14/1972. The appearance of this patent reflects the achievement of a certain technological level, according to which the basic rules for the formation of an aerodynamic configuration were determined for an aircraft with a fixed wing, which provided acceptable aerothermodynamic characteristics in flight modes from the moment of descent from low Earth orbit to landing on the airfield. A technology was created for the use of the latest at that time light heat-shielding materials that withstand aerodynamic heating of the nose of the body and the leading edges of the wing to a temperature of about 1600 ° C. On the basis of these main, as well as a number of other achievements, the design of the rocket-plane spacecraft was created, in which a reusable marching rocket propulsion system of the second stage was placed in the rear part. The rocket fuel for this installation was located in an external disposable block of tanks.

However, at the stage of operation of the Space Shuttle MTKS, miscalculations were fully manifested that were made during the formation of its technical appearance and determined, ultimately, a lower level of efficiency (in terms of the unit cost of putting a payload into a low orbit) compared with the efficiency of transport systems based on disposable launch vehicles. In particular, it became obvious that it was not rational to ensure the return to Earth of a second-stage marching rocket launcher as part of a transport spacecraft. The constant presence on board of this rather heavy installation, but not used in orbital flight, worsens the characteristics of the spacecraft. As you know, in a similar Soviet project, Energia-Buran, which was carried out after the American project Space Shuttle, mid-flight engines of the second stage were no longer placed on the Buran orbital ship, but on the central block of the Energia rocket. At the second stage of the project development (known as Energia-2, or GK-175, 1989), it was planned to combine the central rocket carrier block with a transformable cargo compartment and make this second stage completely reusable, specialized only in putting payloads into a low orbit. In the USA, from the mid-80s, various options for MTKS modernization have also been studied, based on the existing potential of the Space Shuttle system (see the projects Space Shuttle C, Space Shuttle Z, Advanced Launch System, etc.). Apparently, the greatest variety of layouts and proposals for the specialized functions of the MTKS modules in the development of the Space Shuttle system is contained in patent US 4834324 dated 05/30/1989.

Changed views on what should be the rational configuration of a partially reusable transport space system are reflected in patent US 5143327 from 09/01/1992. The core of the rocket launcher arrangement in accordance with this patent is a block of fuel tanks (in the most complete configuration for oxygen, methane and hydrogen) combined in tandem. Four reusable winged modules with a pair of marching liquid rocket engines (LRE) on each are attached to the lower part of this block of tanks in a batch scheme, and a payload that is launched into space is placed on top of the block of tanks, in particular, it can be a reusable winged orbital ship - a spaceplane . In the variant of the most complete set of technical solutions of this patent it is determined that in the aft part of the winged propulsion module there is a hinged mechanism that is designed to move the twin rocket engines from a given working position from under the bottom of the tank block to a given transport position in the cargo compartment in the middle of the winged hull module. The cargo compartment of a reusable propulsion module is closed by a sash (sashes). At the launch vehicle launch stage, the marching liquid propellant rocket engines are extended under the bottom of the fuel tank block and are mechanically connected to the power structure of this bottom by remotely unlockable locks, and are also interfaced via tear-off connectors to the corresponding fuel lines.

It should be noted that in the first independent claims of the patent US 5143327 there is no indication of the use of the articulated mechanism for the movement of marching rocket engines. This means that the initial layout of the winged propulsion module is the configuration with the fixed location of the marching rocket engines, apparently similar to the layout of the Space Shuttle orbiter.

Each of the four winged reusable engine modules is equipped with the necessary set of elements to ensure a planning flight to the landing aerodrome: aerodynamic control bodies, appropriate instrumentation and landing gear. At least two of the four returned engine modules also have means to ensure orbital maneuvering — an orbital maneuvering engine (DOM) and a reactive control system (DCS) with orientation.

In the patent US 5143327 proposed the following method of using the described reusable propulsion modules. At the start of the launch vehicle, marching rocket launchers on all four modules are switched on. At the launch site, when reaching speed M≈3, two diametrically opposite engine modules are separated from the tank block, the rocket engine is switched off on them and transferred to the planning trajectories to the aerodrome of the cosmodrome. Marching rocket launchers of the two remaining engine modules accelerate the rocket system to near orbital speed, after which the rocket engine is also turned off and separated from the tank block. Further, these modules are transferred to the orbital flight path using the orbital maneuvering system, and then at a given distance from the aerodrome of the cosmodrome they give an impulse to exit from orbit. Planning descent in the atmosphere, pre-landing maneuvering, and landing approach for each of these two propulsion modules is planned to be carried out in the same way as it is done on the Space Shuttle. Thus, all four propulsion modules must complete the flight with a non-motor landing at the aerodrome of the cosmodrome in pairs, with a small difference in the moments of landing in the pair and about an hourly delay interval between the pairs. The payload module, after separating it from the disposable tank unit, is informed at a predetermined height of the corresponding impulse to transfer to the support or working orbit, and the tank unit along the ballistic trajectory enters the safe destruction zone in the upper atmosphere.

Evaluation of the totality of technical solutions of the patent US 5143327 from the standpoint of the modern scientific and technical level allows us to indicate a number of insufficiently perfect properties both in the form of reusable propulsion modules, and in the formation of the concept of a launch vehicle as a whole.

Firstly, the development of various MTKS projects with vertical launch of the launch vehicle and horizontal landing of its reusable modules (such systems are designated in the foreign literature by the abbreviation VTOHL - vertical take-of and horizontal landing) showed the preference for separating reusable first-stage blocks at a speed of M≈6 ÷ 7 , but not on M = 3. It is determined that when separating the stages of the launch vehicle at speeds up to M≈7, it is potentially possible on the characteristic flight paths of the returned first stages to dispense with the use of special heat-shielding materials on a significant scale in the design of such stages. Along the way, we can recall that for such aircraft as the MiG-25, SR-71 and MiG-31, the speed M≈3 is included in their operating range. For winged spacecraft designed to launch from orbit with an initial speed of M≈25, the weight of specialized tiled thermal protection is about 9% of the weight of its own spacecraft design (Space Shuttle, Buran). Therefore, space modules, due to the presence of special thermal protection, are overweight for the aviation speed range and therefore it is inefficient to use the same winged propulsion modules as forming a marching propulsion system of both the first and second stages. The separation speed M≈7 is characterized by a different appearance of reusable first stages (see, for example, patents RU 2321526, US 6616092, US 6612522, US 6450452, RU 2148536 and RU 2053936).

Secondly, the elements of relatively short propulsion modules, which protrude significantly beyond the mid-section of the central block of tanks, significantly increase the aerodynamic drag of the carrier at the stage of its flight in dense layers of the atmosphere. It is clear that overcoming this additional resistance will require rocket fuel, while reducing the proportion of the payload weight.

Thirdly, the possibilities for the formation of rocket carriers of various capacities based on the propulsion modules of US Pat. No. 5,143,327 are extremely limited. For example, it is easy to imagine the configuration of a launch vehicle with two winged modules (one module with the power plant function of the first stage, and the second module with the power plant function of the second stage). But it is problematic to form a rocket carrier of six or more modules, since the protruding wings create a serious obstacle to their unification around the central block of fuel tanks.

According to the characteristic parameters of movement during the descent in the atmosphere and landing on the airfield, the winged modules of the second stage of the launch vehicle are similar to reusable planning spacecraft. This circumstance makes it possible to use, to a certain extent, the achievements in aerodynamic layouts and thermal protection of cruise spacecraft for evaluating the technical solutions of salvage propulsion modules of the second stage. Although the target functions of these aircraft belong to different classes, it makes sense to be guided by one general principle in the formation of their appearance with regard to the operations of returning to Earth: to preserve the most important systems, power devices and other equipment with a high unit cost for repeated use. From the standpoint of this principle, for example, it is irrational to save empty fuel tanks or gas cylinders: for serious thermal protection of such voluminous, but thin-walled structures, it will be necessary to spend a noticeable fraction of the mass put into orbit. Apparently, for most applications, it is preferable to accelerate to the orbital speed not the thermal protection of such tanks, but instead to increase the proportion of the payload mass by an appropriate amount.

One of the characteristic reflections of the state of the art in the field under consideration is the evolution of representations in RSC Energia named after S.P. Korolev ”on the appearance of a manned transport spacecraft (see patents RU 2083448 dated 07/10/1997, RU 2220077 dated 12/27/2003 and RU 2334656 dated 09/27/2008). In the last of these patents, an original profiling of the bow of the ship was proposed, which significantly reduces the heat flux to the leading edges of the fixed swept wing. It is alleged that with a spacecraft dimension corresponding to a weight of up to 15 tons, the feasibility of the proposed design on the basis of the existing heat-shielding materials tested at Buran is ensured. The effectiveness of the proposed solution is illustrated by evaluating the heat flux to the leading edges of the wing of a well-known ship of the Space Shuttle or Buran type, but reduced to a size corresponding to a weight of 15 tons (that is, approximately twice as much in linear dimensions). For such a hypothetical object, the maximum heat flux increases approximately twice, and this excludes, according to the authors of the patent, the feasibility of its design based on modern materials.

Based on the data of patent RU 2334656, the feasibility of a number of technical solutions of the earlier patent US 6193187 of February 27, 2001 relating to the layout of the reusable second stage (indicated in the patent as “reusable spacecraft”) and the reusable booster block (indicated in the patent as “ reusable orbit transfercraft "). These devices have folding wings, which are compactly pressed to the body at the stage of launching and fixed in the unfolded working position (approximately parallel to the horizontal construction) at the stage of descent. Given the level of aviation development and the fact that the launch of the second stage is planned to be made from the bow of a hypersonic accelerator aircraft, no one has yet been able to create a real prototype with an unprecedented stability system, the dimensions of such a second stage, as well as an acceleration unit, should fall into the weight class up to 15 tons. The nose of the second stage is made transformable, having a rotating segment in the center. The rotation angle of this movable segment relative to the vertical axis is 180 °. The convex spherical side of the segment is a heat shield at the descent stage, and locks are placed on the opposite surface of this element, which fix the payload module (or an accelerating block with a payload) at the output stage. For another spacecraft from the patent in question - the booster block - the bow is made of four rotary heat-shielding flaps-petals, which in the open position allow compact pairing of this block with the payload, and at the descent stage in the closed position form a hemispherical head fairing that covers payload locks. The options for two technical solutions of US Pat. No. 6,193,187 on dividing the bow of the returned spacecraft into segments determine the not very good position of the boundaries of the sections between these segments (gaps between the wings) - just in the region of maximum temperatures, which is associated with significant problems in providing reliable thermal protection. Special measures are required (about which nothing is written in the patent) to block the penetration of sections of the fairing between the segments of the fairing into the structure of the spacecraft of hot gas. Recall that for all known flying winged spacecraft, nasal coca was made as a single piece, withstanding working heat to a temperature of about 1600 ° C.

It should be noted that technical solutions for providing thermal protection of relatively thin wings are known that were successfully applied in 1982-1984 on the Bor-4 series space planes (the mass of the device in orbit is 1074 kg, the length of the body is 3859 mm, for more details see www.buran. ru ). The wings of these apparatuses are made folding and, in sections of the descent trajectory with the most intense aerodynamic heating, they are turned to such an angle that the direction of the relative flow velocity vectors near the leading edges of the wing is oriented approximately along these edges, and therefore the occurrence of braking zones on the windward surface of the wings was excluded. In addition, an evaporative cooling system for heated surface sections was implemented in the thin wings of the Bor-4 space planes (see patent SU 1840531 of 05.27.2007). In addition to the main results on the development of thermal protection “Burana” on these devices, the effectiveness of aerodynamic control in the roll channel at hypersonic speeds due to the asymmetric deflection of the wing consoles was first confirmed under natural conditions.

In the aggregate of a number of main distinguishing features, the layout decisions of the Bor-4 spaceplane, as well as the wing structure of the head fairing of the upper stage according to US Pat. No. 6,193,187, were selected as prototypes for the present invention application.

A common drawback of most winged spacecraft configurations known to the author is the extensive bottom area in the area of the rear of the hull. The most significant negative impact of this area on the flight performance of space planes is manifested at subsonic speed, and this effect is especially critical on the landing mode.

On the whole, deriving a generalized assessment of the level of technology achieved to date, we have to admit that, despite the abundance of various spacecraft projects and a certain progress in the implementation of some of them, nevertheless, the dominance of disposable launch vehicles and a worthy replacement for a disposable transport spacecraft continues in space exploration. “Union” was never created.

Disclosure of the invention.

Based on the fact that the problem of creating an effective MTKS remains relevant, the synthesis of the transformable layout of the planning spacecraft (spaceplane) as one of the basic elements of such a system is set as a task of the invention, and this ensures the combination of the following properties in a single design:

- the compactness of the launch configuration of the spaceplane should allow it to be coupled with minimal costs with other modules of the rocket system;

- aerothermodynamic characteristics, stability and controllability of the spaceplane in the configuration of hypersonic flight should be no worse than the best analogues of the previous generation;

- The maximum aerodynamic quality of the spaceplane in the landing configuration should be higher than that of analogues of the previous generation.

The selected concept of ensuring the return to Earth of reusable low-orbit reusable modules of the rocket system defines a lot in common in their appearance and design. However, taking into account the specifics of the targeted use of such modules can lead to different versions of technical solutions for individual units, even when the same principle is implemented. In relation to the present application for the invention, the layouts of two spacecraft variants are defined:

- reusable propulsion and instrument module (MDM);

- winged module of a transport spacecraft (KMTKK).

As a result of combining in a tandem scheme one or more MDMs with a block of rocket fuel tanks, the second stage of the launch vehicle is obtained. And as a result of combining the KMTKK in a batch scheme with disposable modules: a conformal inhabited orbital compartment and rocket blocks of the emergency rescue system - orbital maneuvering engines (CAC / HOUSE), we get a transport spacecraft.

Common to the solution of the problem of the invention for the two selected options of spaceplanes is that they can have the same aerodynamic shape in the configurations for the stages of descent and landing, and at the stage of descent with the wing consoles raised, this form should be classified as a “supporting body”, and landing stage - to the "tailless" class. The main target unit (for the MDMM is the forward nozzle / nozzles of the rocket engine, for the KMTKK - the docking unit) is located in the bow of the hull and is closed at the indicated stages of flight by heat-shielding flaps that form the head fairing. The nasal drain is made in the form of a shaped block of heat-resistant material with an upper limit of the working temperature range of at least 1650 ° C (for example, of carbon-carbon composite material) and fixed on the front of the lower shutter of the head fairing. In the closed position, the fairing flaps (at least one lower and two upper) are fastened with tightening locks, providing the specified conditions for thermal sealing at the interface between these moving elements together, as well as with the middle part of the spacecraft hull. The tail of the hull is tapering, mating with minimal aerodynamic losses with the elevator. The space console wing consoles should be controlled, rotating relative to the axes, which are approximately parallel to the longitudinal axis of the device. If necessary, a cooling system, for example, an evaporative type, should be installed in the wing consoles. In the starting configuration, the consoles are turned to the housing to the extreme preset position (raised), ensuring the compactness of the module. In addition to aerodynamic controls on space planes, rocket control devices — DCS and HOUSES — should also be installed. The information and control complex should provide control corresponding to the flight program at all its stages - from start to stop after runway runway.

In the initial MISP configuration, the head fairing flaps should be open and compactly displaced to the middle part of the spacecraft hull so as not to interfere with the normal operation of the mid-flight rocket launcher. The nodes of the power interface with the block of fuel tanks of the second stage, as well as the nodes of interface with the fuel lines are located behind the midsection section of the MPDM housing mainly on its upper part.

The flaps of the head fairing KMTKK should be open at the stage of orbital flight during the operation of docking with another space object, and on the inner surface of the flaps can be installed devices that control the docking. The kinematic diagram of the movement of the upper flaps of the head fairing and the placement of the workplace of at least one pilot in the nose of the spaceplane are performed so that in the open position between the flaps there is a sufficient viewing sector in the front hemisphere, allowing the pilot to visually control the docking process. At the flight stages, the “take-off” and “descent” in the CMTC passengers are placed on the lodgements in the pose “lying on their back” mainly perpendicular to the plane of symmetry of the space plane, and the axis of rotation of each of the lodges is also predominantly perpendicular to the same plane and offset from the longitudinal axis of the passenger’s body (in indicated position) closer to his chest.

The set of technical solutions formed in this application for the invention is focused on obtaining the following technical results:

- expanding the reachability area of landing aerodromes and lowering the tension of the final landing regimes — leveling and maintaining — in comparison with analogs (due to the higher maximum aerodynamic quality of the new space plan layout);

- providing the possibility of forming, on the basis of unified MDPM configurations of the second stages of the launch vehicle with a multiple-varying power marching rocket launcher;

- ensuring the minimum acceptable volume in the KMTKK cabin for a comfortable position of passengers in the modes of departure and descent relative to the variable vector of the overload acting on them;

- providing the most natural visualization to the pilot of the external space at his workplace at the KMTKK during docking in orbit, as well as during the descent and landing modes, which improves the reliability of these operations.

Thus, in accordance with the claimed invention, a positive effect is expected due to the fact that when forming space plan layouts and their integration into the MTKS design, a new combination of the following distinctive features should be added:

a) One of the main units of the spaceplane (LRE for MDM or docking unit for KMTKK) is located in the bow of the hull so that it is possible to operate normally with open shutters of the head fairing and reliable heat protection during the descent and landing stages when the shutters are closed, when they form a head fairing.

b) The nose heat-resistant coke should be made as an element of a single heat-shielding structure as a part of the lower wing of the head fairing of the spaceplane.

c) The aft part of the spacecraft's hull must be made tapering to the elevator and have a shape determined by the conditions for minimizing its contribution to the aerodynamic drag of this aircraft in the subsonic range of flight speeds.

d) On the tapering stern of the MPDM hull, mainly from above, there should be its counter-coupled units in tandem with the stern of the rocket fuel tank unit of the second stage of the launch vehicle.

e) The upper flaps of the head fairing of the manned spacecraft in the open position should be fixed so that in the resulting viewing sector of the front hemisphere, the pilot can visually control the approach to the cooperating object during the docking operation in orbit.

f) Passengers of the spaceplane at the stages of launching, launching and landing should be placed on self-orienting lodgements in the “lying on back” pose, with the axis of rotation of each of the lodgements should be approximately perpendicular to the plane of symmetry of the spacecraft and offset from the longitudinal axis of the passenger’s body to the area of his chest.

g) On the upper part of the manned spacecraft’s hull, mainly on its sides, there should be located interface nodes in a packet scheme with SAS / DOM rocket blocks and in a wedge-shaped scheme with a conformal orbital compartment.

h) When controlling the rate of dissipation of mechanical energy in the landing mode, a simultaneous coordinated rotation of the wing consoles in opposite directions (that is, their simultaneous raising or lowering), which changes the aerodynamic quality of the spaceplane, can be used.

Brief Description of the Drawings

The main technical solutions of the present application for the invention are illustrated in ten figures.

Figure 1. The layout of the second stage of the launch vehicle with one single-engine MPM.

Figure 2. Starting configuration of a single-engine MPM.

Figure 3. A variant of the hinge mechanism for moving the lower leaf of the MDFM fairing.

Figure 4. MISM intermediate configuration with closed head fairing flaps and with folded wing consoles.

Figure 5. MISM in the configuration of descent in the atmosphere.

6. MDPM in landing configuration - wing consoles lowered and landing gear released.

7. The manned spacecraft in the launch configuration as part of the MTKS head part put into orbit.

Fig. 8. The orbital configuration of a partially reusable transport spacecraft.

Fig.9. Manned spacecraft in the configuration of descent in the atmosphere at hypersonic speed.

Figure 10. Typical application pattern for MPPM.

The following notation is used in these figures:

1 - spacecraft body

2 - LPRE nozzle

3 - block rocket fuel tanks of the second stage

4 - space warhead

5 - aft fairing of the second stage - hood MDM

6 - upper (central) power element of the interface of the second stage with the MPM

7 - lateral power element of the interface of the second stage with the MPM

8 - upper heat shield of the head fairing

9 - lower heat shield of the head fairing

10 - rotary wing console

11 - central aerodynamic steering wheel

12 - side node interface with the power element 7

13 - nozzle RSU course channel

14 - the upper nozzle of the DCS of pitch / roll channels

15 - preferred placement points on the trailing edge of the heat shield

mates of clamps

16 - preferred placement of pull locks

17 - interceptor

18 - hinged mechanism for moving the lower sash of the MPPM

19 - the Central node of the power interface MDPM with the block of tanks of the second stage

20 - fixation unit of the wing console in the starting configuration

21 - elevon

22 - a breakaway interface connector with the fuel line

23 - nozzle of the orbital maneuvering engine

24 - container brake parachute landing

25 - the main landing gear

26 - front wheel landing gear

27 - module of the first stage of the launch vehicle

28 - section of the trajectory of the launch vehicle to the point of separation of steps

29 - section of the flight path of the second stage of the launch vehicle

30 - reference orbit

31 - section of the flight path of the module of the first stage after the moment of separation

32 - the trajectory of the flight MDM to the aerodrome after leaving orbit

33 - airfield runway

34 - heat resistant nasal coca

35 - inhabited spacecraft compartment

36 - space pilot

37 - space crew member seat

38 - self-orienting tool tray of a spaceplane passenger

39 - docking unit

40 - access hatch between the spaceplane and the orbital module

41 - conformal orbital module

42 - inhabited compartment of the orbital module

43 - engine block DCS orbital module

44 - aggregate compartment of the orbital module

45 - instrument-aggregate compartment of the spaceplane

46 - passageway cover

47 - border of the working area of the self-orienting tool tray

48 - missile blocks SAS / HOUSE

49 - interface unit missile block SAS / HOUSE with a spaceplane

50 - spacecraft interface unit with a conformal orbital compartment

51 - dashboard of the crew commander

52 - cockpit glazing

53 - interface unit of the orbital compartment with the block of fuel tanks of the second stage

54 - the point of separation of the modules of the first stage of the launch vehicle from the Central unit

55 - the point of separation of the MPM and the head from the block of tanks of the second stage of the rocket

56 - permissible nozzle rocket block SAS / HOUSE

The implementation of the invention

In the previous sections of the present application for the invention, the significance of two types of winged cowl-winged space planes for the formation of the second generation MTKS second stage layouts was disclosed. Figure 1 in isometric projection shows the second stage of the MTKS with single-engine MPM in the starting configuration with the open nozzle 2 of the main missile launcher (see distinctive feature a). The aft part of the body 1 of the MDPM is connected in an opposite manner in tandem with the aft part of the tank block 3 of the second stage (see hallmark d) by means of the power coupling elements 6, 7 and the corresponding locks. A space head part 4 is attached to the other end of the tank block 3. A part of the MDPM hull with the wing consoles folded and the wing fairing flaps displaced to the extreme open position is closed by the aft cowl 5 of the second stage. This fairing should provide not only acceptable aerodynamic characteristics of the stern of the second stage, taking into account the creation of favorable conditions for the functioning of the nozzle 2 of the mid-flight rocket launcher, but also shield the MISP from the influence of other adverse gas-dynamic factors on it at the stage of elimination. In appearance, the layout of the second stage of the launch vehicle with the MISP may not differ much from the typical layout of the disposable second stage, except perhaps for a slightly longer length.

Figure 2 shows in more detail an isometric projection of the initial configuration of the MDPM. Above the tapering stern of the hull 1, the “hut” has pivoting wing consoles 10. The upper heat-shielding flaps 8 of the head fairing are compactly pulled from above and fixed on top of them. To improve the compactness of the initial MISP configuration, the trailing edges of the flaps 8, as well as the lower flap 9, are made curly, with protrusions in the middle and cutouts at the edges. The middle protrusions on these edges should be reciprocal fragments for closing the niches on the MDPM case in the area where the LPRE nozzle begins. The nasal parts of the valves in the open position are lowered into these niches. Cropped edges allow the sash to be closer to the body in the open position. In Fig. 2 it is clearly seen that the head fairing flaps offset by a sufficient distance back allow us to provide normal operating conditions for the nozzle 2 of the second stage sustainer rocket launcher directed relative to the casing forward (see distinctive feature a). This figure also shows the preferred locations for the DCS nozzles for control in the yaw channel (key 13) and in pitch / roll channels (key 14). The counterparts of the latches 15 on the wings are located near the extreme points of their convex surfaces, so that in the open position of the wings it would be convenient to fix them on the corresponding elements of the stern of the hull or wing consoles. The locations of the tightening locks 16 for fixing the flaps in the closed position should then be determined on the mating longitudinal protrusions of the housing surface.

Figure 3 shows one of the simplest possible schemes of the hinge mechanism 18 for moving the lower sash 8 of the head fairing. The dotted line indicates the position of the structural elements in the closed position, and the solid line in the open position. The above diagram illustrates one of the key ideas of the solution to ensure the compactness of the MISM starting configuration, namely, the use of the characteristic narrowing of the rocket engine contours in the area of the nozzle neck to shift the relatively bulky nose of the lower heat-shielding shutter into the formed “hollow” - a heat-resistant cocoon 34. It should It should be noted that in the MISM arrangement with two nozzles, the conditions for “immersion” of nasal coca between them are better than in the case of a single nozzle.

Figure 4 shows the intermediate configuration of the MDPM when the heat-shielding flaps 8, 9 are moved forward and secured by tightening locks 16 in the form of a head fairing covering the nozzle. The interface between the lower cusp and the upper cusps is determined taking into account the conditions of rational integration into the design of heat-resistant coca 34 (see hallmark b), as well as obtaining a compact starting configuration of the MPPM. Thermal protection at the borders of the fairing flaps can be provided on the basis of proven technical solutions used in the designs of the first generation aerospace vehicles, for example, in the designs of the flaps of the Burana landing gear niches. In order to create sufficient forces for a tight fit of the edges, providing the required operating conditions for heat-shielding seals, the tightening locks 16 should be placed not only at the points marked in Fig. 4, but also in other places around the perimeter of the wings. In the form of an open dog (position 19), the central unit for interfacing the MPM with the power element 6 is indicated, through which the main part of the thrust of the main missile launcher is transmitted to the tank unit 3 (see distinguishing feature d). The places of approach of the lateral power elements 7 to the MDPM body are indicated in the form of grooves 12 on the root of the wing console 10. These grooves automatically overlap when the wing consoles rotate from the starting to the working position during the descent and landing stages. In the starting position, to ensure structural rigidity, the ends of the wing consoles 10 are connected by a lock 20. Elevons 21 can be added to the set of MDPM aerodynamic controls (consisting of the elevator 11 and rotary wing consoles 10) as well as synchronous multidirectional deviation with elevator) perform the functions of the air brake.

To illustrate the effect on the layout of the spacecraft of the distinctive feature c, Fig. 5 shows the MDM in the view "from above from above" from the rear hemisphere in the configuration of hypersonic flight. The wing consoles 10 are rotated to the operating position by an angle corresponding to the self-balancing mode (approximately 45 ÷ 50 ° upwards relative to the horizontal construction of the body). In addition to the previously described elements, this figure shows the preferred locations of the following units: nozzles of the orbital maneuvering engine 23, explosive connectors 22 of the fuel lines, as well as the container 24 of the landing brake parachute. In a selected perspective, an interceptor 17 is shown on the upper surface of the wing console. Elevons 21 and / or interceptors 17 can be included in aerodynamic control bodies of the space plan if, for any reason, the control efficiency is via rotary wing consoles 10 and elevator 11 will be insufficient.

A typical view of the landing configuration MDPM shown in Fig.6. Chassis racks 25, 26 are released, wing consoles 10 are lowered almost to the horizontal of construction. This configuration is aimed at obtaining close to the maximum possible bearing properties of the layout and acceptable characteristics of the directional stability of the apparatus.

Inscribed in the contours of a typical shape of the space head part 4, the layout of the transport spacecraft with a manned spaceplane in the launch configuration is presented in Fig.7. The movable flaps 8, 9 of the head fairing close the docking unit 39, located in the bow of the hull of the spaceplane 1 and oriented along the longitudinal axis of inertia of the entire assembly. The middle part of the spacecraft’s hull is occupied by the inhabited compartment 35, in which the workplace of the crew commander 36 is arranged in accordance with the aviation canons. This relates to the orientation of the pilot's posture in the seat 37 and the location of the information-control field - the dashboard 51, glazing 52, as well as the controls (not shown). Passengers of the spaceplane at the stages of launching, launching and landing are placed on self-orienting lodges 38 in the “lying on their back” pose (see hallmark f). This allows, firstly, to rationally use the features of the shape of the hull narrowing from the middle elliptical to the stern wedge-shaped with a horizontal rib and, secondly, to improve the conditions of overload tolerance by adjusting the orientation of the passenger’s body to the line of action in the "chest-back" direction. Between the inhabited compartment 35 of the spaceplane and the inhabited compartment 42 of the conformal orbital module 41 there is a passage through the sealed hatch 40. The module 41 is connected to the spacecraft 1 by the locks 50 (see hallmark g). Power elements 53 at the aft end of the module 41 are used for its detachable connection with the block of tanks 3. At the same end can be installed the engine block RSU 43, designed primarily to control the orientation of the spacecraft at the stage of orbital flight. Missile blocks 48, attached by locks 49 to the spacecraft’s hull 1 (see hallmark g), can be used as a propulsion device in the emergency rescue system (CAC) and as an orbital maneuvering engine (HOUSE). In the first of the listed applications, after opening the locks 50, a sufficient number of blocks 48 must be switched on at the same time to remove the spaceplane from the emergency launch vehicle. The conditional point of application of the resultant from the jet rods of these blocks should be in front of the center of mass of the separated package. The second use case is focused on the separate inclusion of one or more blocks 48, when for the orbital maneuver of the spacecraft it is required to give it a sufficiently large momentum. The nozzles 56 of blocks 48, after entering orbit, must be rearranged so that the direction of action of the thrust passes through the center of inertia of the spacecraft.

On Fig shows the main configuration of the transport spacecraft for the stage of orbital flight. Flaps 8 and 9 of the head fairing are open. The pilot 36 can visually control part of the space in the front hemisphere through the viewing sector between the open upper wings 8 (see hallmark e). The cover 46 of the access hatch 40 between the inhabited compartments of the spaceplane and the conformal orbital compartment is open. In the instrument-aggregate compartment 45 of the spaceplane, mainly sophisticated equipment (electronic components, electromechanical systems, etc.) is located, and in the aggregate compartment 44 of the orbital module, mainly containers with stocks of liquid and gaseous working bodies.

The configuration of a manned spacecraft depicted in FIG. 9 at the launching stage is similar in state to the main aerodynamic elements of the MISM configuration in FIG. 5. A comparison of the positions of passengers in Fig.7 and Fig.9 allows to illustrate the effect of the application of the hallmark f. Self-orienting lodges 38 provide the orientation of the bodies of passengers in a position that minimizes the negative impact on them of overload. In fact, the orientation with respect to the overload vector of the “chest-back” direction for the passengers of the spaceplane at the stages of launching and launching is approximately the same as, for example, for the cosmonauts of the Soyuz at the corresponding stages of flight. Since the overall level of overloads on the spaceplane on the “planning” descent trajectories is lower than on the “sliding” or “ballistic” trajectories of the Soyuz capsule, therefore, the level of comfort for passengers on this indicator in the new means of descent from orbit should be higher.

The sequence of basic operations for the use of MISP, typical of one of the two layout options for space planes with a wing head fairing, is shown schematically in FIG. 10. The key differences in this sequence from the closest analogue in the method of application (see US Pat. No. 5,143,327) consist in operations to move the head fairing flaps, as well as in the particularities of the variants of the laws of formation of control actions in the landing mode using the wing console layout options.

A typical MPLM flight program as part of the MTKS according to the claimed invention should include the following sequence of operations:

1. Carry out a vertical launch of the launch vehicle, which then flies along the launch path 28 to the separation point 54. Figure 10 shows the launch vehicle in a three-block arrangement, consisting of two reusable modules of the first stage 27 and the second stage in the configuration of figure 1 with one MDM.

2. After dividing at point 54 at a speed of approximately M≈7 and turning off the marching rocket engines on reusable modules 27 of the first stage, they plan a flight in automatic mode to the landing aerodrome (s) along the corresponding paths 31. The second stage consisting of modules 1, 3 , 4 continues the flight along the trajectory of the output 29 to the point of separation 55.

3. On the flight path 29 to point 55, when the specified suborbital speed is reached, the marching rocket engine on the MPD is turned off.

4. After the separation of the second-stage modules at point 55, the MPDM and the payload module are removed from the tank block 3. The MPDMs from the starting configuration (see FIG. 2) are converted to the descent configuration in the atmosphere (see FIG. 5).

5. The payload module 4, which can be either a disposable spacecraft or KMTKK, is brought into a given reference orbit according to the standard procedure. MDPM with the help of DCS (pos.13, 14) is oriented in space as needed, then at the appropriate moment turn on the HOUSE (pos.23) for a specified time in order to continue the flight along the orbital path 30.

6. At a predetermined distance from the selected landing aerodrome (according to the regular flight program, this should be the aerodrome of the cosmodrome), the MPPM descend from the orbit and perform the DCS (pos. 13, 14) and DOM (pos. 23) for this purpose.

7. The MDPM flight path is controlled by aerodynamic surfaces on the hypersonic section of the descent in the same way as the Bor-4 series was controlled - the asymmetric rotation of the wing consoles 10 relative to a given balancing angle of their layout. In the longitudinal channel, it is additionally possible to use a rudder of height 11, which at large angles of attack acts as a balancing shield.

8. Aerodynamic control on the portion 32 of the spacecraft flight path at moderate supersonic and subsonic speeds is carried out:

i. in the pitch channel - due to the deviation of the elevator (pos. 11);

ii. in the roll channel - due to the asymmetric rotation of the wing consoles (10);

iii. in the track channel - due to the coordinated deviation of the elevator (pos. 11) and asymmetric rotation of the wing consoles (pos. 10);

iv. in the speed channel - due to the deviation of the elevator (pos. 11) and / or the symmetrical raising / lowering of the wing consoles (pos. 10).

9. In the final section of the alignment trajectory, the angle of the transverse V of the wing consoles 10 is reduced to a minimum predetermined value (apparently, to about -5 °). The landing planes landing control on the runway (pos. 33), as well as run control, is carried out in accordance with well-known aviation procedures.

The following considerations are the basis for determining the composition of the executive bodies for the speed control channel. As is known, (see, for example, Guskov Yu.P., Zagainov GI Flight control of aircraft. M: Mashinostroenie, 1980, p. 166), when forming aircraft automatic landing control systems, the function of regulating flight speed is assigned primarily to the subsystem propulsion traction control. At the same time, faster aerodynamic rudders also make a certain contribution to ensuring the high quality of the speed control process, simultaneously being used in other control loops. On reusable first-generation vehicles (Space Shuttle, Buran), an air brake was used as an effective means of solving a number of problems associated with obtaining acceptable aerodynamic characteristics, which was made in the form of two drop-down halves of the rudder. For speed control purposes, an air brake on the “Buran” was used at M≤0.8 (see the reusable orbiter “Buran”. Yu.P. Semenov, G.E. Lozino-Lozinsky et al .; - M .: “Mechanical Engineering” 1995 , p. 300). According to the effect obtained, speed control using an air brake on these devices has become the equivalent of control actions provided by changing the thrust of an air propulsion device in airplanes.

Regulation of aerodynamic drag by means of an air brake leads to corresponding changes in the aerodynamic quality coefficient of the aircraft, affecting the determination of the established values of the planning angle and speed. But aerodynamic quality can also be controlled by changing the load-bearing properties of the aircraft, which just happens with the simultaneous raising / lowering of the wing consoles (since the projection of the effective bearing surface on the horizontal building plane of the device changes). Thus, in the proposed version of the spaceplane layout, the set of governing bodies “rotary wing consoles + elevator” is for the control system a functional analogue of the “air brake + elevons” set previously used on Space Shuttle and Buran devices. Or (somewhat more distant) the aggregate “air mover + elevator (elevons)”, typical for most aircraft with an automatic approach system.

It is assumed that for space planes with a new layout, the reference program of speed change during approach will be determined mainly by the layout angle (the angle of the transverse V) of the wing consoles, and the program will be such that the reserves along the angle of the elevator in both directions are nominally approximately the same, with compliance with typical restrictions on landing on other flight parameters. These principles, in fact, determine how speed control should be implemented.

The main differences in the method of using the spaceplane as part of a transport spacecraft (compared with the above method for MIS) are as follows:

I. In the starting configuration, the heat-shielding flaps 8, 9 are closed and form the head fairing.

II. In the event of an emergency during the launch of the MTKS, it is possible to urgently separate the spaceplane from the emergency rocket carrier and the orbital compartment at any point on launch path 28, 29, followed by landing on the runway.

III. Flaps of the head fairing 8, 9 necessarily open only if you need to use the docking unit 39.

IV. The habitable spacecraft compartment, closed by hatch 40, can be used as a lock chamber when the astronaut enters outer space through the hatch of docking unit 39.

V. In addition to closing the head fairing flaps 8, 9 (if they were open), before leaving the orbit, the spaceplane and the conformal orbital compartment 41 are undocked, and after the descent pulse has been worked out, the hull 48 of all SAS / DOM missile units is also undocked.

As is known, in the Bor-4 experimental apparatuses, the main purpose of which was to conduct full-scale tests of the Buran thermal protection elements, only two asymmetrically deflected wing consoles in the transverse channel were used as aerodynamic controls. Nevertheless, a control system with such truncated capabilities ensured the solution of the problems of modeling the hypersonic section of the Burana flight path and bringing the Bora-4 to the specified splashdown area. In the present application for the invention, it is proposed to expand to approximately 130 ÷ 150 ° the total range of angles of rotation of the console of the wing of the spaceplane in comparison with the known previous counterparts (projects "Spiral", "Bor-4", MAKS). Such a solution makes it possible to ensure the compactness of the launch plan of the space plan and to maximize its load-bearing properties at supersonic and subsonic planning speeds. The addition of the third moving surface — the elevator — to the aerodynamic controls of the spaceplane creates the possibility of full regulation (separately and combined) through four channels, namely, pitch, roll, track and airspeed. The positive effect of the proposed technical solution for the configuration of aerodynamic controls is not only that in the heat-stressed design of space planes it is possible to halve the number of movable steering surfaces (from six, like the Space Shuttle, or seven, like the Buran) three. It is also positive that the opportunity arises to rationally get rid of the difficulties with ensuring stability and controllability in certain ranges of flight speeds that were identified in the Shuttle assembly.

In conclusion of this section, it should be noted that for the implementation of the proposed technical solutions in the layout and design of the spaceplane, the technological level at which Space Shuttle and Buran were created is potentially sufficient. But the use of a new combination of these solutions should allow achieving a higher level of flight-technical and operational characteristics of the next generation of MTKS than it was possible to obtain at the previous one.

Claims (10)

1. Planning spacecraft with a wing head fairing, comprising a heat-bearing supporting body, rotary wing consoles, the axis of articulation of which with the body are approximately parallel to its longitudinal axis, aerodynamic and jet steering wheels, instrumentation to provide control of orbital flight, descent and landing of this device on airfield, retractable landing gear (landing gear), detachable units of power, pneumohydraulic, electrical and information interface with mates the rocket fuel tank unit, as well as the module of the propulsion carrier propulsion system, characterized in that the nozzle of the propulsion module of the propulsion system located in front of the planning spacecraft body is oriented forward relative to this body, the bow of which also contains one lower and at least two upper heat-shielding movable flaps of the head fairing, compactly superimposed in the launch configuration on the middle part of the body of the planning spacecraft so as to provide the standard operation of the marching power plant, and having the possibility of their advance forward with the formation in the landing configuration of the surface of the head fairing around the nozzle of the marching power plant, while the adjacent wings are pairwise tightly connected to each other and to the edge of the middle part of the hull of the planning spacecraft by locking devices, creating the necessary conditions for the functioning of the heat seal along the edges of these flaps, and the lower cusp of the head fairing contains nasal heat a good cook, made as an element of a single heat-shielding design, in addition, the aft part of the planning spacecraft’s hull is made tapering and paired with the elevator contours, while the counter-coupling of the aft parts of the planning spacecraft and the rocket fuel tank block of the second stage of the launch vehicle is made according to the tandem scheme, moreover, detachable interface nodes are located mainly in the upper part and on the sides of the tapering stern compartment of the planning spacecraft. 2. Planning spacecraft with a wing head fairing according to claim 1, characterized in that the marching propulsion module contains two nozzles oriented forward relative to the hull of this planning spacecraft. 3. Planning spacecraft with a flap head fairing according to any one of claims 1 and 2, characterized in that the trailing edges of the heat-protective flaps of the head fairing are shaped so that in the closed position they cover the niches in the front of the hull that are designed to be located in them front segments of these valves in the open position. 4. Planning spacecraft with a wing head fairing according to any one of claims 1 to 3, characterized in that on the upper surface of the wing consoles are spoilers. 5. Planning spacecraft with a wing head fairing according to any one of claims 1 to 4, characterized in that the elevons are located on the wing consoles. 6. Planning spacecraft with a wing head fairing, comprising a heat-bearing supporting body, rotary wing consoles rotating relative to axes approximately parallel to the longitudinal axis of its body, aerodynamic rudders, jet engines and instrumentation to provide control of orbital flight, descent and landing of this device on aerodrome, retractable landing device (landing gear), characterized in that the planning spacecraft contains an inhabited cargo-passenger compartment with a working m by the astronaut’s pilot’s nose, interfaced with a detachable conformal orbital compartment through an airtight hatch, detachable components of power, electric, pneumohydraulic and information interfaces, which are located mainly on the upper part of the planning spacecraft’s body, a docking unit installed in the bow of the body of this device and tightly closed by movable heat-shielding flaps of the head fairing in the areas of removal, descent and landing with the possibility of opening the flaps n the orbital portion of the flight so that conditions are provided for direct visual control from the workplace of the pilot-cosmonaut of a given approach sector with a cooperating object, the nose heat-resistant cock is made as an element of the unified heat-shielding structure of the lower wing of the head fairing, and the aft part of the body of the planning spacecraft is made tapering and paired with elevator contours, mainly on the sides of the upper surface of the hull of the planning spacecraft, power units are installed new detachable interface in a batch scheme with orbital maneuvering engine blocks, the movable nozzles of which in the starting position ensure the operation of the emergency rescue system, and lodgement units are located in the inhabited area of the cargo-passenger compartment. 7. Planning spacecraft with a wing head fairing according to claim 6, characterized in that self-orienting passenger lodgements are installed in the habitable zone of the cargo-passenger compartment of the hull, adjusting the position of the passenger bodies so that at the flight stages the "launch" and "descent" overload affects them in the "chest-back" direction, the axis of rotation of each of the lodgements is predominantly perpendicular to the plane of symmetry of the planning spacecraft, and passengers are placed on the lodgements in the "supine" also approximately perpendicular to this plane of symmetry. 8. Planning spacecraft with a wing head fairing according to any one of claims 6 and 7, characterized in that on the upper surface of the wing consoles there are spoilers. 9. Planning spacecraft with a wing head fairing according to any one of claims 6 to 8, characterized in that elevons are located on the wing consoles. 10. A method for controlling the return from orbit of a planning spacecraft, which has a layout according to any one of claims 1 to 9, including performing orbital operations of orientation, descent from orbit and stabilization, ensuring the entry of this spacecraft into the planet’s atmosphere with specified parameters, corrective actions on the orientation of the planning spacecraft and the trajectory of its movement with the help of aerodynamic and gas-dynamic rudders in accordance with the given law of controlling the trajectory of descent and landing on the flight strip of the aerodrome, characterized in that the control actions using aerodynamic forces are carried out in the pitch channel by deflecting the elevator, in the roll channel by asymmetric deviation of the wing consoles, in the track channel due to the combined deviation of the elevator and asymmetric deviation of the wing consoles, and in As an additional means of control actions in the airspeed channel, synchronous raising or lowering of the wing consoles is used.

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