WO2013039426A1 - Gliding spacecraft with folding nose fairing (variants) and method for controlling the return thereof to the landing field - Google Patents

Abstract

The inventions relate to partially reusable launch and space transportation vehicles, and are intended to improve the configuration of such craft, as well as to a method for controlling the return of the reusable modules thereof to the landing field. The gliding spacecraft (space plane) comprises a thermally-protected lifting body faired with an elevator, wing panels which can be extended in the transverse plane, gas dynamic members for controlling the orientation and flight path of the space plane, instrumentation for providing controlled descent and landing, as well as aircraft landing gear. The nose fairing of the body is made up of movable thermally-protected panels, namely a bottom panel with a heat-resistant nose cone and (nominally) two top panels. If a reusable propulsion and instrumentation module of a second stage launch vehicle is converted into a space plane, then the panels cover the nozzle (or two nozzles) of the main rocket assembly in the reentry configuration. If the space plane is a reusable module from a space transportation vehicle, then the nose fairing panels cover the docking assembly in the insertion and reentry stages. When a reusable propulsion and instrumentation module is in the launch configuration, the wing panels are tented over the upper part of the body of the space plane, and the displaced top panels of the nose fairing lie compactly on them. The bottom panel of the fairing is also displaced and lies compactly against the underside of the middle part of the body of the space plane. The main technical results of the inventions are an increase in the aerodynamic properties of the two chosen types of space plane in the aircraft portion of flight, and the possibility of creating a series of second stage launch vehicle designs based on standardized reusable propulsion and instrumentation modules, in which the total power of the main rocket assembly can vary greatly.

Description

 PLANNING SPACE VEHICLE WITH HINGED HEAD CIRCULATOR (OPTIONS) AND METHOD OF CONTROL ITS RETURN-

COMMUNICATION AT THE AERODROME The invention relates to the field of rocket and space technology: reusable space transport vehicles and launch vehicles, and is aimed at improving the layout of such devices, as well as the way to control the return of reusable modules to the airport.

 The patent US 3702688 from 14.1.1.1972 (IPC B 64c 37/02) “SPACE SHUTTLE VEHICLE AND SYSTEM”, which describes the aerodynamic layout of a two-stage reusable transport space system (MTKS), is known. On the first (atmospheric) and second (space) reusable steps, it was planned to establish fixed wings. The tail and marching rocket launchers were located in the aft of the hulls. These steps were also equipped with a landing gear and other necessary equipment for landing on the airfield. One of the main distinguishing features of the construction of the second stage was the use of passive heat protection from light heat-insulating materials that can withstand aerodynamic heating of the nose of the body and the leading edges of the wing to temperatures of about 1600 ° C. In the middle part of the body of the second stage there was a cargo compartment for useful load, and in the bow - the crew’s cabin.

 The decision to combine in a single design of the second (space) stage the provision of functions: a transport spacecraft for the crew, a cargo container for delivery into orbit and return from orbit of the payload, the means of returning a reusable marching rocket launcher is ineffective when there is no need to return to Earth loads comparable in weight to those delivered into orbit. The implementation of such a solution led to the creation of a structure whose own weight was approximately twice the maximum weight of the useful load delivered into orbit. The gain from the reusability of the second stage did not block the loss from about a three-fold decrease in the transport efficiency of the launch vehicle as a whole.

In the Energy-Buran space system (www.buran.ru), which was later created by Space Shuttle, the second stage marching engines were no longer placed on the Buran reusable orbiter, but on the central block of the Energy rocket, which provided greater versatility of the missile system. However, the marching rocket propulsion system of the second stage was not saved.

 In US Pat. No. 5,143,327 of September 1, 1992, IPC B 64 G 1/40, the core of the launch vehicle configuration 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 the payload that is launched into space is placed on top of the block of tanks, in particular, it can be multiple Calling winged orbital ship - spaceplane. In the aft part of each winged propulsion module, there is a hinge mechanism that is designed to move the twin rocket engines from a predetermined working position from under the bottom of the tank block to a predetermined transport position in the cargo compartment in the middle of the wing module housing. (However, the initial layout of the winged propulsion module is a configuration with the fixed location of the main propellant rocket engines in the rear of the module housing.) Each of the four winged reusable propulsion modules is equipped with the necessary set of elements for ensuring the planning flight to the landing aerodrome: aerodynamic organs control, appropriate instrumentation and landing gear. At least two of the four returned engine modules have means to ensure orbital maneuvering — an orbital maneuvering engine (DOM) and a reactive orientation control system (DCS).

 The disadvantages of patent US 5143327 is:

 Firstly, 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.

 Secondly, the elements of the propulsion modules that are significantly protruding beyond the mid-section of the central block of tanks increase the aerodynamic drag of the carrier at the stage of its flight in dense layers of the atmosphere. To overcome the resistance, additional rocket fuel is required, while the payload is reduced.

Thirdly, the formation of rocket carriers of various capacities, for example, of six or more modules, is problematic, since stationary protruding roofs Leahs pose a serious obstacle to the integration of such modules around the central block of fuel tanks.

 The development of various MTKS projects with vertical launch of the launch vehicle and horizontal landing of its reusable modules (such systems are indicated in 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, and not on M = 3. It was determined that when separating the stages of the carrier rocket at speeds up to M ~ 7, it is potentially possible on the characteristic flight paths of the returning first stages to dispense with the use of special heat-shielding materials on a significant scale in the design of such stages. For winged spacecraft intended for descent from orbit with an initial velocity of M ~ 25, the weight of a specialized tiled thermal protection is about 9% of the weight of its own spacecraft design (Space Shuttle, Buran). Consequently, space modules, due to the presence of special thermal protection, are overweight for the airspeed range and therefore it is ineffective to use the same winged propulsion modules as forming a marching power plant, both of the first and second stages.

 In a number of patents, RSC Energia named after S.P. Koroleva ”(RU 2083448 dated 10.07.1997, IPC B 64 G 1/62, RU 2220077 12/27/2003, IPC B 64 G 1/12 and RU 2334656 09/27/2008, IPC B 64 G 1 / 14) reflects the evolution over the past decade and a half of the ideas of the developers of that corporation regarding the appearance of a reusable manned transport spacecraft - from a capsule with a supporting hull to a spaceplane with an original bow, which became known as the Clipper (www.buran.ru).

The aerodynamic layout of “Clipper” refers to the type of “tailless” with keels at the ends of the wing. In the center, a disposable adapter compartment with a docking unit is connected to the aft part of the body of its reusable module. Jobs pilots "Clipper" are in the bow. Such a layout solution, traditional for aviation, provides the convenience of controlling the spaceplane at the landing mode, but does not create conditions for direct visual control of the docking process in space. Although instrumental means of dock control have long been mastered in astronautics, nevertheless, the possibility of direct visual control would increase the reliability of this important orbital operation. In patent RU 2220077, a spacecraft layout was proposed in which the crew created We have the possibility of direct visual control of the processes of rapprochement and docking. But for this version of the spacecraft, the supporting body and the parachute landing system were chosen. In the transition to the winged layout of the spacecraft, preference was given to the convenience and reliability of landing on the airfield due to the fact that instrumental information support was left for docking operations in space.

 Also known patent US 6193187 from 02.27.2001, IPC B 64 G 1/14, which presents technical solutions related to the layout of the reusable second stage and reusable booster block (essentially the third stage). These devices have folding wings, which are compactly pressed to the body during the withdrawal stage and are fixed in the unfolded working position (at a given angle relative to the construction horizontal) during the descent stage. The wing consoles have elevons and vertical wingtips (keels) with aerodynamic rudders. 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 flat surface of this element, which fix the payload module (or an accelerating block with a payload) at the output stage. In another spacecraft from the patent in question - the booster block - the bow is made of four rotary heat-shielding petal flaps, which in the open position allow compact conjugation of this block with the payload, and form a hemispherical at the descent stage in the closed position head fairing covering payload locks.

The disadvantages of the two technical solutions of US Pat. No. 6,193,187 on dividing the nose of the spacecraft returned into segments are the unsuccessful positioning of the boundaries between these segments, partially passing on the lower surface through the maximum temperature range. This makes thermal protection in such places very vulnerable, since it is not possible to completely get rid of the presence of gaps between the wings. Special measures are needed to block the penetration of hot gas into the spacecraft structure through the interfaces between the fairing segments. 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. The wing consoles have only two positions and cannot be used in control, which is carried out using five aerodynamic surfaces.

 Known technical solutions for providing thermal protection for relatively thin wings, applied in 1982 - 1984 on spaceplanes of the Bor-4 series (the mass of the device in orbit is 1074 kg, the length of the body is 3859 mm, www.buran.aO. These devices have wings In sections of the descent trajectory with the most intense aerodynamic heating, the cantilevers were rotated so 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 the surface of the wings of the braking zones In addition to the main results of testing the Buran thermal protection on these devices, the aerodynamic control in the roll channel at hypersonic speeds was confirmed for the first time under natural conditions due to asymmetric deflection of the wing consoles. the layout of this apparatus of other aerodynamic controls did not allow to obtain full control in such modes as landing.

 A common drawback of most of the known configurations of winged spacecraft is the vast bottom region in the area of the rear of the hull. The most significant negative effect of this area on the flight performance of space plans is manifested in the subsonic speed range, and such an effect on the landing mode is especially critical.

In US Pat. No. 5,143,327 of September 1, 1992, IPC B 64 G 1/40, the following method is proposed for using the described reusable propulsion modules. At the launch of the launch vehicle, marching rocket launchers on all four modules are switched on. At the exit site, when reaching a speed of M ~ 3, two diametrically opposed engine modules are separated from the tank block, the mid-range rocket engines are switched off and transferred to the planning path to the aerodrome of the cosmodrome. Marching rocket launchers of the two remaining propulsion modules accelerate the rocket system to near orbital speed, after which they also turn off the LRE and separate them from the tank block. The twin rocket engines are moved to the luggage compartment of the spaceplane and the compartment is closed with sashes. Then, these modules are transferred using the orbital maneuvering system to the trajectory of the orbital flight, and then at a given distance from the aerodrome of the cosmodrome they give an impulse to exit from orbit. The planning descent in the atmosphere, the pre-landing maneuvering and the landing approach for each of these two propulsion modules were planned to be carried out in the same way as it was done on the Space Shuttle. Thus, all four propulsion modules must complete the flight with a non-motor landing at the aerodrome.

 There is also a known method of controlling the return from spacecraft orbit according to patent US 6193187 dated 02.27.2001, IPC B 64 G 1/14, in which, before entering the atmosphere, heat-shielding flaps are shifted and fixed in the form of a head fairing, orbital orientation operations are performed, the descent from the orbit and stabilization, which provide the spacecraft entering the planet’s atmosphere with the given parameters, carry out corrective actions on the spaceplane using aerodynamic and gas-dynamic rudders in accordance with the given law of descent trajectory control landing on the runway airport.

 The disadvantage of these methods is that at least four control surfaces and drives for their movement are involved in their implementation, which is associated with a complication of the design, an increase in weight, and a decrease in control reliability.

 The objective of the invention is to provide a space plan of a transformable layout with a compact launch configuration and the possibility of aerodynamically improved interfacing with other modules of the rocket system while maintaining good aerodynamic characteristics, stability and controllability of the space plane in hypersonic flight configuration and landing modes.

 The technical result consists in increasing the maximum aerodynamic quality of the spaceplane in the landing configuration, shifting the mating boundaries of the head fairing flaps to lower temperatures, reducing the aerodynamic resistance of the aft part of the spacecraft’s hull, reducing the number of movable steering surfaces in the basic configuration, and, as a result , reducing the weight of the spaceplane. Space planes in the MDIS version improve the capabilities of modular construction of a space rocket system by varying the number of marching rocket engines at the second stage of the launch vehicle. The space plan in the KMTKK layout at all stages of the flight created conditions for a direct visual review of the front hemisphere sector during piloting, as well as improved conditions for the portability of passenger overloads.

The technical result is achieved by the fact that the planning spacecraft (space plane) with a wing head fairing contains 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 rudders, boron equipment for providing control of the orbital flight, launching and landing of the spaceplane at the aerodrome, retractable landing gear (landing gear), detachable units of power, pneumohydraulic, electrical and information interfaces with mating parts on the rocket fuel tank unit, as well as the marching power module rocket launchers, oriented forward relative to the housing, one or two nozzles located in its nose, containing one lower and at least two upper heat-shielding movable flaps g a space fairing compactly superimposed on the middle part of the spacecraft so as to enable the regular operation of the marching propulsion system, and they can be advanced and rotated forward with the formation of the configuration of the head fairing around the nozzles of the marching propulsion system. the flaps are pairwise tightly connected to each other and to the edge of the middle part of the spacecraft hull by locking devices, creating the necessary conditions for the functioning of the thermal seal along the edges of these a cooler, and the lower wing of the nose fairing contains a heat-resistant nose cock made as an element of a single heat-shielding construction; moreover, the aft part of the spaceplane body is made narrowing in the plane of symmetry and smoothly interfaced with the elevator contours with the smallest permissible losses in aerodynamic quality, In this case, the counter-coupling of the aft parts of the spaceplane and the block of rocket fuel tanks of the second stage of the carrier rocket is made according to the tandem scheme, with detachable conjugation nodes located predominantly tween the top and on the sides of the tapered aft compartment space plane.

 The technical result is also achieved by the fact that the trailing edges of the flaps of the head fairing of the planning spacecraft are made curly and, in the closed position, cover the body niches designed to locate the leading edges of the flaps in the open position.

 The technical result is also achieved by the fact that at the planning spacecraft (spaceplane) with a wing head fairing on the upper surface of the wing consoles there are interceptors.

 The technical result is also achieved by the fact that the planning spacecraft (spaceplane) with a wing head on the wing consoles have elevons.

The technical result is achieved by the fact that the planning spacecraft (spaceplane) with a wing head fairing comprises a heat-bearing supporting body, rotary wing consoles, rotating relative to the axes approximately parallel part of the longitudinal axis of the spacecraft’s hull, aerodynamic rudders, jet engines and instrumentation to control the orbital flight, the launch and landing of the spaceplane at the aerodrome, a retractable landing gear (landing gear), an inhabited cargo-passenger compartment with the workplace of the cosmonaut nose, which mates with a detachable conformal orbital compartment through an airtight hatch, detachable nodes of power, electric, pneumohydraulic and information interfaces, which are located mainly on the upper part of the spacecraft’s hull, a docking unit installed in the bow of the spacecraft’s hull, tightly closed by movable heat-shielding flaps of the head fairing at the sites of launch, descent and landing, with the possibility of their opening on the orbital section of the flight to enable direct visual control from the workplace of a pilot-cosmonaut of a given approach sector with a cooperating object, the nose heat-resistant coke is made as an element of a single heat-protective design and the lower wing of the head fairing, and the aft part of the hull of the space planes is made tapering in the plane of symmetry and smoothly interfaced with the contours of the elevator with the least permissible losses in aerodynamic quality, mainly on the sides and on the upper part of the hull of the space planes are mounted nodes of detachable coupling for batch scheme with orbital maneuvering propulsion units, the movable nozzles of which in the starting position ensure the operation of the emergency rescue system, and in the cargo-passenger zone the spacecraft’s battered compartment, the nodes of the lodgement hitch are located.

 The technical result is also achieved by the fact that in the planning spacecraft (spaceplane) with a wing head fairing in the passenger-and-freight zone of the inhabited compartment, self-orienting lodgements are installed for placing passengers mainly in the "lying on the back" position along the rotation axis and facing it, and the rotation axis each lodgement is approximately perpendicular to the plane of symmetry of the spaceplane, and the center of inertia of the lodgement, together with the passenger, is between the axis of rotation and the back of the lodgement.

 The technical result is also achieved by the fact that the planning spacecraft (spaceplane) with a wing head fairing contains interceptors on the upper surface of the wing consoles.

The technical result is also achieved by the fact that at the planning spacecraft (spaceplane) with a wing head fairing, elevons are located on the wing consoles. In addition, the technical result is achieved by the fact that when controlling the return from the orbit of the space plane, including performing orbital operations of orientation, descent from the orbit and stabilization, which ensure that the space plane enters the planet’s atmosphere with specified parameters, and that corrective actions are applied to the space plane using aerodynamic and gasdynamic rudders in accordance with the law governing the descent and landing on the runway of the aerodrome, control actions using the aerodrome dynamical forces produced in the channel due to the combined yaw deflection wing panels and the elevator in the roll channel by asymmetrical deflection wing panels, and as an additional means of control actions in the channel airspeed use a synchronous lifting or lowering of the wing panels

The main technical solutions of this invention application are illustrated in ten figures.

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

MDPM.

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

 FIG. 3. A variant of the hinge mechanism for moving the lower shutter of the MDFM fairing.

 FIG. 4. Intermediate configuration of the MDPM with closed head fairing flaps and with folded wing consoles.

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

 FIG. 6. MDPM in landing configuration - wing consoles are lowered and the landing gear is released.

 FIG. 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. The manned spacecraft in the configuration of descent in the atmosphere at hypersonic speed.

 FIG. 10. Typical application pattern for MPPM.

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 to lead to different versions of technical solutions for individual units even when implementing the same principle.

 This application for an invention defines the layout of two spacecraft variants:

 - reusable propulsion and instrument module (MDM);

 - winged module of a transport spacecraft (KMTK).

 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 according to the KMTKK packet scheme with disposable modules: a conformable inhabited orbital compartment and rocket blocks of the emergency rescue system (CAC) - orbital maneuvering engines (DOM), we get a transport spacecraft.

 A common thing in solving the problem of the invention for the two selected options of space planes is that they can have the same aerodynamic shape in configurations for the stages of descent and landing.

 One of the main units of the spacecraft (LRE for MDLM 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 head fairing. In FIG. 1 shows in isometric projection the second stage of the MTKS with a single-engine MPMM in the starting configuration with an open nozzle 2 of a marching rocket launcher. The aft part of the MPPM hull 1 is connected in tandem with the aft part of the tank block 3 of the second stage by means of power coupling elements 6 , 7 and associated locks. A space head part 4 is attached to the other end of the tank unit 3. A part of the MISM body with folded wing consoles and the head fairing flaps displaced to the extreme open position is closed by the aft fairing 5 of the second stage. The aft fairing provides not only acceptable aerodynamic characteristics of the aft second stage taking into account the creation of favorable conditions for the functioning of the nozzle 2 of the main missile launcher, but also shields the MISP from the influence of other unfavorable gas-dynamic factors on it during the removal stage.

In FIG. 2 shows in more detail an isometric view of the MISP startup configuration. The hinged wings of the wings 10 are folded over the tapering stern of the hull 1 by the “hut”. upper heat-proof flaps 8 of the head fairing. 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 at these edges should be reciprocal fragments to close the niches on the MDPM body in the area where the LPRE nozzle begins. The bow parts of the trunks in the open position are lowered into these niches. Cropped edges on the trailing 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 shifted a sufficient distance back allow us to ensure the normal operating conditions of the nozzle 2 of the second stage marching rocket launcher directed forward relative to the casing. This figure also shows the preferred locations of the DCS nozzles for control in the yaw channel (position 13) and in pitch / roll channels (position 14). The counterparts of the latches 75 on the wings are located near the extreme points of their convex surfaces, so that in the open position the wings would be conveniently fixed to the corresponding elements of the stern of the hull or wing consoles. The location of the tightening locks 16 for fixing the flaps in the closed position is made on the reciprocal longitudinal protrusions of the surface of the housing.

 In FIG. 3 shows one of the simplest possible schemes of the hinge mechanism 18 for moving the lower wing 8 of the head fairing. The dotted line indicates the position of 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 throat to shift the relatively bulky nose of the lower heat-shielding shutter into the resulting "heat-sink" - heat-resistant cocoa 34. B A MISP arrangement with two nozzles provides better conditions for “immersion” of nasal coca between them than in the case of a single nozzle.

In FIG. 4 shows the intermediate configuration of the MDM, when the heat-shielding flaps 8, 9 are moved forward and secured by pull-up locks 16 in the form of a head fairing covering the nozzle. The interface between the lower flap and the upper flaps was determined taking into account the conditions of rational integration into the design of heat-resistant coca 34, made as an element of a single heat-shielding structure as a part of the lower flap of the head fairing of the spaceplane, as well as obtaining a compact MISP launch configuration. The boundary pairing of the valves is located on the upper surface of the spaceplane. Thermal protection at the borders of the fairing flaps may be liver based on technical solutions used in the construction of aerospace equipment of the first generation, for example, in the design of the flaps of the niches of the Burana landing gear. In order to create sufficient forces for a snug fit of the edges, providing the required working conditions for heat-shielding seals, the tightening locks 16 are placed not only in those indicated in FIG. 4 points, 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 rocket launcher is transmitted to the block of tanks 3. 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 holes automatically overlap when the wing consoles rotate from the starting position to the working position during the descent and landing stages. In the starting position, to ensure rigidity of the structure, 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 a rudder 11 and rotary wing consoles 10), as a means of longitudinal-transverse control, and also (with synchronous multidirectional deviation with the elevator) performing the functions of an air brake.

 The aft part of the spacecraft’s hull is made tapering in the direction of the elevator, and has a shape determined by the conditions for minimizing its contribution to the aerodynamic drag of the aircraft in the subsonic range of flight speeds. In FIG. Figure 5 shows the MDM in the “side to top” view from the rear hemisphere in the configuration of hypersonic flight. The wing consoles 10 are rotated to the operating position at 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 the selected flight, the interceptor / 7 is shown on the upper surface of the wing console. Elevons 21 and / or interceptors 17 can be included in the aerodynamic control bodies of the space plan if, for any reason, the control efficiency in the basic configuration by rotary consoles of the wing 10 and the elevator 11 will be insufficient.

A typical view of the MPDM landing configuration is shown in FIG. 6. The landing gear 25, 26 are released, the wing consoles 10 are lowered almost to the horizontal level. This configuration is focused on getting close to the maximum possible the essential properties of the layout and acceptable characteristics of the track stability of the device.

The layout of the transport spacecraft with a manned spaceplane inscribed in the contours of a typical shape of the space head part 4 in the launch configuration is shown in FIG. 7. Movable flaps 8, 9 of the head fairing close the docking unit 39, located in the bow of the spacecraft 1 and oriented along the longitudinal axis of inertia of the entire assembly. The middle part of the space plan’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 refers to the orientation of the pilot's posture in the seat 37 and the location of the information-control field - the dashboard 57, glazing 52, as well as the controls (not shown). Passengers of the spacecraft at the stages of launching, launching and landing are placed on self-orienting beds 38 in the “lying on their back” position, and the axis of rotation of each lodgment should be approximately perpendicular to the plane of symmetry of the spacecraft and offset from the inertia axis of the passenger’s body to the area of his chest , that is, the center of inertia of the lodgement, together with the passenger, is between the axis of rotation and the back of the lodgement. This allows, firstly, to rationally use the features of the shape of the body narrowing from the middle ellipsoidal to the feed wedge-shaped with a horizontal rib, and, secondly, to improve the conditions of overload tolerance by adjusting the orientation of the passenger body to the line of action in the direction "Chest - back." Between the inhabited compartment 35 of the space plan and the inhabited compartment 42 of the conformal orbital module 41 there is a passage through an airtight hatch 40. On the upper part of the manned spacecraft’s hull, mainly on its sides, there are interface units in the package design with the SAS / DOM rocket blocks and in a wedge-shaped pattern with a conformal orbital compartment. The module 41 is connected to the spacecraft’s hull 1 by locks 50. The power elements 53 on the aft end of the module 41 are used for its detachable connection with the block of tanks 3. At the same end, an RSU 43 engine block can be installed, designed primarily for control orientation of the spacecraft at the stage of orbital flight. Missile units 48, attached by locks 49 to the spacecraft’s hull 1, are used as a propulsion device in the emergency rescue system (CAC) and as an orbital maneuvering engine (HOA). In the first of these applications in an emergency situation, after opening the locks 50, at the same time, a sufficient number of blocks 48 are turned on to remove the spaceplane from the 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 separated package. The second use case is focused on the separate inclusion of one or several 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 into orbit, are rearranged so that the direction of action of the thrust passes through the center of inertia of the spacecraft.

 In FIG. Figure 8 shows the basic configuration of a transport spacecraft for the orbital flight phase. Flaps 8 and 9 of the head fairing are open and allotted by an articulated mechanism to the sides of the housing. Pilot 36 can visually control part of the space in the front hemisphere through the viewing sector between the open upper leaves 8 and visually control the approach to the cooperating object during the docking operation in orbit. 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-and-assembly compartment of the 45 spaceplane, mainly sophisticated equipment (electronic components, electromechanical systems, etc.) is located, and in the aggregate compartment of the 44 orbital module, it is mainly containers with reserves of liquid and gaseous working bodies.

 Depicted in FIG. 9, the configuration of the manned spacecraft at the launching stage is similar in state to the main aerodynamic elements of the MISM configuration in FIG. 5. A comparison of the passenger positions in FIG. 7 and FIG. 9 allows one to illustrate the effect of applying the self-orientation of lodgements. Self-orienting elements 38 provide the orientation of the bodies of passengers in a position that minimizes the negative impact of overload on them. In fact, the orientation with respect to the overload vector of the “chest - back” direction for the passengers of the space plan at the stages of launch and descent 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 “moving” or “ballistic” trajectories of the Soyuz capsule, therefore, the level of comfort for passengers according to this indicator on the new means of descent from orbit should be above.

The sequence of basic operations for using the MISP, typical of one of the two options for layout of 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 consist in operations to move the head fairing flaps, as well as in the features of the variants of formation laws control actions at the landing mode using the layout options of the wing consoles in combination with the deviation of the elevator.

 The 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 path of launch 28 to separation point 54. In FIG. 10 shows a carrier rocket in a three-block configuration, consisting of two reusable modules of the first stage 27 and the second stage in the configuration of FIG. 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 carry out a planned flight in automatic mode to the landing aerodrome (s) along the appropriate paths 31. The second stage consists of modules /, 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 liquid propellant rocket engine is switched off.

 4. After the separation of the second-stage modules at point 55, the MPMM and the payload module are removed from the tank block 3. The MPMM from the starting configuration (Fig. 2) is converted into a descent configuration in the atmosphere (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. MDMs using DCS (pos. 13, 14) are oriented in space as needed, then at the appropriate moment they 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, it should be the aerodrome of the cosmodrome), the MPPM orbit will be removed from the orbit, reapplying DCS (pos. 13, 14) and HOUSE (pos. 23).

 7. The MDPM flight path is controlled by aerodynamic surfaces on the hypersonic descent section by 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 over section 32 of the spacecraft flight path at moderate supersonic and subsonic speeds is carried out: in the pitch channel — due to the elevator deflection (pos. 11); in the roll channel - due to asymmetric rotation of the wing consoles (pos. 10); in the track channel - due to the coordinated deviation of the elevator (pos. 11) and asymmetric rotation of the wing consoles (pos. 10); in the speed channel - due to the deviation of the elevator (pos. 11) and / or the symmetrical raising and lowering of the wing consoles (pos. 10).

 9. At 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 (to about -5 °). The runway landing control on the runway (pos. 33), as well as run control, is carried out in accordance with well-known aviation procedures.

 The main differences in the method of using the KMTKK space plan as a part of a transport spacecraft (as compared to the above-described method for MDS) are as follows:

 I. In the starting configuration, the heat shields 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 are required to be opened only when the use of the docking unit 39 is necessary.

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

 V. In addition to closing the shutters 8, 9 of the head fairing (if they were open), before leaving the orbit, the spaceplane and the conformal orbital compartment 41 are undocked, and after working off the descent pulse, the bodies 48 of all SAS / HOUSE.

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 deflectors in the transverse channel of the wing console 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 reduction area. In the present application for the invention proposed to expand to approximately 130 ° -450 ° the total range of rotation angles of the wing of the spaceplane compared with smaller ranges of the known previous analogues (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 use its load-bearing properties at supersonic and subsonic planning speeds as much as possible. 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) over four channels, namely, pitch, roll, yaw 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 spacecraft, it is possible to halve the number of movable steering surfaces (from six, like the Space Shuttle, or seven, like “Buran”) up to 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, which were identified in the Shuttle layout.

 Thus, the possibility of obtaining the claimed technical result in the implementation of the totality of the solutions proposed in this application is determined by the fact that:

 In the aerodynamic layout of space planes, both in the MDM and the CMTC variants, there is no vast bottom region. Compared with analogues and prototypes that have bottom regions, this allows one to noticeably improve such basic aerodynamic characteristics as drag coefficient (reduce it by about 15–25%) and increase the maximum aerodynamic quality by about 1–2 units.

The design of the lower cusp of the head fairing is formed so that its interface with the upper cusps is located on the upper surface of the space plane, where the temperature fields are much smaller (up to several hundred degrees Celsius) compared with comparable areas on the lower surface. This improves the conditions in which the thermal insulation of the joints functions, and, ultimately, improves the reliability of thermal protection.

The number of aerodynamic controls in the basic configuration of spaceplanes has been reduced to a minimum of two rotary wing consoles and elevator, and the working range of angles of rotation of the wing consoles has been significantly expanded. The compact configuration of the MPDM launch vehicle creates the conditions for the use in the second stage of the launch vehicle of dense placement of sets of two variants of MPDM layouts, and the total number of LREs in the kit can be from 1 to 8-10.

In the KMTKK layout, the docking unit is located in the bow of the space plan’s hull; therefore, its working area is in the direct visual field of view from the regular workplace of the pilot. The hinged mechanism for opening the upper flaps is designed so that the flaps, when opened, move to a position on the sides of the hull and do not overlap the pilot viewing sector in the front hemisphere.

KMTKK passengers at the stages of launching into space and returning to Earth are placed on self-orienting lodgements in the inhabited compartment in the “lying on their back” position, and the lodgements are set so that the head – pelvis direction is perpendicular to the plane of symmetry of the spaceplane. In this arrangement, the most significant overloads will affect passengers along the “chest - back” line, which corresponds to an approximate optimal body orientation for tolerance of significant overloads.

 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 space plan, 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 to achieve 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

Claim 1. Planning spacecraft (spaceplane) with a wing head fairing containing a heat-bearing supporting body, rotary wing consoles, the axis of articulated joints with the body being approximately parallel to its longitudinal axis, aerodynamic and jet rudders, instrumentation to control orbital flight, descent and landing of the spacecraft at the airfield, retractable landing gear (landing gear), detachable nodes of power, pneumohydraulic, electrical and information interface with response h asthma on the block of rocket fuel tanks, as well as the module of the propellant marching propulsion system, characterized in that the spaceplane contains one or two nozzles oriented forward relative to the housing, located in its nose, containing one lower and at least two upper heat-shielding movable shutters of the head fairings, compactly superimposed on the middle part of the spacecraft hull so as to enable the regular functioning of the marching power plant and have the ability to extend and rotate forward with the formation of the configuration of the head fairing around the nozzles of the marching propulsion system, while adjacent flaps are pairwise tightly connected to each other and to the edge of the middle part of the space planes by locking devices, creating the necessary conditions for the functioning of thermal sealing along the edges of these wings, and the lower the nose fairing flap contains a heat-resistant nose cock made as an element of a single heat-shielding structure, in addition, the aft part of the space planes body is made tapering in the sim plane metric and smoothly interfaced with elevator contours with the smallest permissible losses in aerodynamic quality, while the counter-coupling of the aft parts of the spaceplane and the rocket fuel tank block of the second stage of the launch vehicle is carried out in tandem, with detachable interface nodes located mainly in the upper parts and sides of the tapering stern compartment of the spaceplane. 2. Planning spacecraft (spaceplane) with a wing head fairing according to claim 1, characterized in that the trailing edges of the head fairing cusps are closed and in the closed position they cover the body niches made for the front edges of the cusps to be located in the open position. 3. Planning spacecraft (spaceplane) with a wing head fairing according to Claim 1 or Clause 2, characterized in that there are spoilers on the upper surface of the wing consoles.  4. Planning spacecraft (spaceplane) with a wing head fairing according to claim 1 or claim 2, characterized in that eleons are located on the wing consoles.  5. Planning spacecraft (space plane) with a wing head fairing containing a heat-bearing supporting body, rotary wing consoles rotating relative to the axes approximately parallel to the longitudinal axis of the spacecraft body, aerodynamic rudders, rocket engines and instrumentation to provide control of orbital flight, descent and landing of the spaceplane at the aerodrome, retractable landing device (landing gear) characterized in that the spaceplane contains an inhabited cargo-passenger compartment with a working place astronaut pilot in the bow, which mates with a detachable conformal orbital compartment through an airtight hatch, detachable nodes of power, electric, pneumohydraulic and information interfaces, which are located mainly on the upper part of the spacecraft’s hull, a docking unit installed in the bow of the spacecraft’s hull, tightly closed by movable heat-shielding flaps of the head fairing in the areas of launch, descent and landing, with the possibility of opening the flaps in the orbital section flight to enable direct visual control from the workplace of the pilot-cosmonaut of a given approach sector with a co-ordinated object, the nose heat-resistant cock is made as an element of a single heat-shielding structure of the lower wing of the head fairing, and the aft part of the space planes body is made narrowing in the plane of symmetry and smoothly It is connected with elevator contours with the least permissible losses in aerodynamic quality, mainly on the sides of the upper part of the spacecraft’s hull. nodes of detachable power interface according to the batch scheme with engine blocks of orbital maneuvering, movable nozzles of which in the starting position provide the emergency rescue system, and nodes of hitching elements are located in the passenger and passenger area of the inhabited spacecraft compartment. 6. Planning spacecraft (spaceplane) with a wing head fairing according to claim 5, characterized in that in the cargo-passenger zone of the inhabited compartment the spaceplane has self-orienting lodgements for placement along the axis of rotation of the lodgement of passengers, mainly in the "lying on the back" position, with the axis of rotation of each lodgement being approximately perpendicular to the plane of symmetry of the spaceplane, and the center of inertia of the lodgement with the passenger is between the axis of rotation and the back of the lodgement.  7. Planning spacecraft (spaceplane) with a wing head fairing according to Claim 5 or Clause 6, characterized in that there are spoilers on the upper surface of the wing consoles.  8. Planning spacecraft (spaceplane) with a wing head fairing according to Claim 5 or Clause 6, characterized in that Elevons are located on the wing consoles.  9. A method for controlling the return from orbit of a space plane, including performing orbital operations of orientation, descent from orbit and stabilization, which ensure that the space plane enters the planet’s atmosphere with predetermined parameters, carries out corrective actions on the space plane using aerodynamic and gas-dynamic rudders in accordance with the specified the law of controlling the path of descent and landing on the runway of the aerodrome, characterized in that the control actions using aerodynamic forces carried out dissolved in yaw channel due to combined deflection and wing panel of the elevator, in the roll channel by asymmetrical deflection wing panels, and as an additional means of control actions in the channel airspeed use simultaneous lifting or lowering of the wing panel. 21 SUBSTITUTE SHEET (RULE 26)