WO2017116256A1 - Aircraft - Google Patents

Abstract

The claimed invention relates to aviation, and more particularly to tailless aircraft. In one possible embodiment, the claimed invention has a fuselage, a wing, and two turbojet engines disposed one above the other in a single nacelle in the plane of symmetry of the aircraft. The fuselage is rigidly fastened to the underside of the nacelle. The wing consists of a centre section, a left panel and a right panel. The centre section of the wing is pivotally fastened to the upperside of the nacelle and is designed such that the angle of incidence of the wing can be adjusted (in the longitudinal plane) relative to the fuselage and the nacelle. Air intakes of the turbojet engines are situated on the underside of the centre section of the wing and on the upperside of the fuselage. The left and right panels of the wing are pivotally fastened to the centre section of the wing and are designed such that their angle of sweep can be adjusted. The wing has a high-lift device.

Description

 O P I S A N I E I Z O B R E T E N I

Aircraft

The invention relates to aircraft (LA) and relates in particular to aircraft.

State of the art

From (http://ru.wikipedia.org/wiki/Vought_F-8_Crusader, [1]), the F-8 Crusader fighter of the American company Vought is known as a high-wing aircraft with a “normal” aerodynamic design. F-8 Crusader was the only serial aircraft in the history of aviation with a variable in-flight wing installation angle (in the longitudinal plane, relative to the fuselage). In take-off and landing flight modes, the wing of the aircraft was set at an angle of 10 °, and in cruise flight mode, the wing of the plane was set at an angle of 1 °.

 Advantages of the F-8 Crusader aircraft: the use of a wing with a variable installation angle allows takeoff and landing with almost the horizontal position of the fuselage, which significantly reduces the required height of the landing gear (and, therefore, reduces the weight of the landing gear and fuselage); Good visibility from the cockpit.

From (Ruzhitsky EI American vertical take-off aircraft. M: Astrel ACT, 2000, p. 177-H 90, [2]) the American experimental aircraft XC-142A vertical take-off and landing “normal Noah "aerodynamic design. In a given airplane, during vertical take-off and landing, the wing is installed (in the longitudinal plane, relative to the fuselage) in a vertical position (rotated through an angle of 100 °) by means of double synchronized screw jacks, and in cruising flight the wing occupies a horizontal position.

In the 6th ⁽ 80s of the 20th century) supersonic military aircraft of the “normal” aerodynamic design quite widely used a wing with variable sweep. This allowed the aircraft, on the one hand, to have good takeoff and landing characteristics (by reducing the sweep of the wing it increased the coefficient of the maximum lift of the wing (Su.max), and, on the other hand, to reduce the aerodynamic drag during flight at maximum supersonic speed (by increasing the sweep of the wing).

 A typical aircraft with variable sweep of the wing is an American F-111 fighter (Bauer P. Aircraft of non-traditional designs. M: Mir, 1991, pp. 130-34, [3]). In F-l l 1, the left and right wing consoles have direct sweep (variable, depending on the flight mode of the aircraft).

 From (US Pat. No. 3,529,790 A, B64C 3/40, published September 22, 1970, [4]), an airplane of a “normal” aerodynamic design is known, in which the wing is configured to change the sweep angle and the angle of installation (attack). In this aircraft, the rotary wing assembly allows you to change both the angle of sweep of the wing and the angle of installation of the wing. However, due to the design of this unit (which is structurally complex), the range of change in the angle of installation of the wing is small (within 0- ^ 5 °).

Most modern military aircraft use brake parachutes to reduce the mileage during landing. Rye can reduce the mean free path when landing on the ZSN-35%. Advantage of brake parachutes: the braking force does not depend on the condition of the runway. (Zaitsev V.N., Rudakov

B. L. Design and strength of aircraft. Kiev. Vishka school. 1978,

C.399, [5]).

 Disadvantages of brake parachutes: each time they have to be dropped from the plane and quickly removed from the runway; they wear out rather quickly ([5], p.402).

 For example, in a Soviet MiG-29 fighter plane, the wing area is 38 m 2, and the area of the parachute canopy is 17 m 2 (http: // www. Airwar.ru/enc/fighter/mig29.html, [6]).

 As you know, planes without horizontal tail (aerodrome “tailless” and “flying wing”), theoretically, can have higher aerodynamic quality (due to lower aerodynamic drag - due to the smaller surface area washed by the incident air flow), in comparison with airplanes of the “normal” aerodynamic design (having horizontal plumage (GO) and vertical plumage (VO)).

 For example, the total relative area of plumage (GO + VO) for the aircraft of the “normal” aerodynamic scheme is 53% of the wing area (Sheinin V.M. Weighted design and efficiency of passenger aircraft. M: Mashinostroenie, 1984, p.194, [7]).

From (Sobolev D. A. Airplanes of special schemes. M .: Mashinostroenie, 1989, p. 61, [8]), an experienced American subsonic bomber XB-35 aerodynamic scheme "flying wing" (the first flight in 1946), which the maximum lifting force coefficient of the Su.max wing was equal to 1.5 (for comparison, another American B-29 subsonic bomber (the first flight in 1942) had “normal maximum aerodynamic design, the coefficient of the maximum lifting force of the wing Su.max was equal to 2.3). As a result, to maintain the same landing speed on the XB-35, it was necessary to increase the wing area. This reduced the value of the maximum aerodynamic quality of the XB-35 aircraft.

 For the KhV-35 aircraft, the maximum aerodynamic quality was 22.6 ([8], p. 63, table 1.5).

 For comparison, another American subsonic bomber of a “normal” aerodynamic design with a V-52 swept wing (first flight in 1952) has a maximum aerodynamic quality of 21.5 (DA Sobolev. Centennial history of the “flying wing”. M: RUSAVIA, 1998, p.208, [9]).

 As can be seen from the foregoing, both the KhV-35 bomber (performed according to the “flying wing” aerodynamic scheme) and the B-52 bomber (performed according to the “normal” aerodynamic scheme) have close values of maximum aerodynamic quality.

One more example. The “normal” aerodynamic “Mirage” FIF airplane of the French company Dassault has a wing load of 6.2 kilonewtons per square meter (kN / m ² ), and the similar “tailless” aerodynamic “Mirage” aerodynamic aerodynamic airplane “Mirage” the firm’s wing load is 3.5 kN / m. ”Moreover, the landing speeds of both aircraft are 235 kilometers per hour (IK Kostenko. Flying wings. M .: Mashinostroenie, 1988, p. 14. [10 ]).

As can be seen from the foregoing, the tailless aerodynamic design (and the flying wing aerodynamic design), in practice, cannot realize its theoretical advantages in terms of increasing the maximum aerodynamic quality compared to the normal aerodynamic design. This is hindered by the only (still unsolved) problem: in the aerodynamic design “tailless” (and “flying wing”) it is impossible to use highly efficient take-off and landing wing mechanization (as in aircraft of the “normal” aerodynamic design).

This is explained by the following. When using highly effective wing take-off and landing mechanization (for example, when the flaps or flaps are deflected), the wing center of pressure is shifted back, which leads to an increase in the diving moment on the wing (Petrov KP Aerodynamics of elements of aircraft. M.: Mechanical Engineering, 1985, p.1 13, [1 1]).

 On an airplane without horizontal tail (on the “tailless” and “flying wing”) this dive moment has nothing to balance.

 This forces the tailless scheme (and the “flying wing”) on the plane to increase the landing speed and take-off and landing distance within the required limits, to increase the wing area, which increases the aerodynamic drag, and, therefore, reduces the value maximum aerodynamic quality of the aircraft in cruise flight, and does not allow to use all the theoretical advantages of the aircraft of this aerodynamic scheme.

 Thus, the aforementioned drawback of the tailless aircraft (and the “flying wing”) aerodynamic scheme nullifies its theoretical advantages compared to the “normal” aero dynamic aircraft.

Closest to the claimed invention is a tailless aircraft of the aerodynamic scheme known from (RU 2486105, B64C 30/00, published June 27, 2013, ([12]). In the version of the supersonic airplane (ΦΗΓ.12ί13) it has , fuselage, wing, two turbojet (TRD) engines placed in a common nacelle one above the other in plane of symmetry of the aircraft. The nacelle is attached to the wing from its underside. The fuselage is attached to the wing by means of a common engine nacelle. In supersonic engine air intakes, a vertical multistage wedge of constant geometry common to both engines was used as a generator of shock waves. The engine intake channels are separated by a horizontal partition. Thus, in this aircraft, the engine intake channels are, on the one hand, located on the lower side of the wing, and on the other hand, located on the upper side of the fuselage. At a supersonic flight speed, the shock waves from the vertical wedge and the air intake shell of the engine nacelle sit only on the lower surface of the wing, which reduces the angle of attack of the wing, and, therefore, reduces the aerodynamic drag of the wing and improves the aerodynamic quality of the aircraft.

Advantages of the prototype: the location of two engines on top of each other in the plane of symmetry of the aircraft allows you to have a minimum total midship of the fuselage and engines, which increases the aerodynamic quality of the aircraft; the adopted arrangement of engines makes it possible to refuse vertical plumage, since in case of failure of one of the engines there is no destabilizing (unrolling) moment along the course, which increases the aerodynamic quality of the aircraft; The position of engine air intakes on the upper side of the fuselage prevents foreign objects from entering them when the aircraft moves on the ground. four

 7

 Disclosure of invention

The task of the invention is to improve the flight performance of the prototype.

 Obviously, if such a problem can be solved, then this is a “non-obvious” solution for a specialist who is knowledgeable in the corresponding field of technology, since it has not been solved for the prototype and other known analogues.

The invention, in one of the possible variants of its execution, has the following essential features in common with the prototype: the aircraft of the aerodynamic scheme "tailless", has, fuselage, wing, at least one device made to create power rods, for instance, ^'jet engine (WFD) disposed in the plane of symmetry of the aircraft, the shroud of the above WFD is located on the lower side of the wing and the fuselage with the top side.

 Distinctive features from the prototype are: the wing consists of a center wing, left and right consoles, the wing center section is made with the possibility of changing its installation angle (in the longitudinal plane) with respect to the fuselage, the left and right wing sections are pivotally attached to the wing center section and are made with the possibility of changing the angle of their sweep, the wing has take-off and landing mechanization.

By changing the angle of installation of the wing of the claimed invention, compared with the prototype [12], it is possible: to maintain the horizontal position of the fuselage and the engine nacelle of the WFD (in any flight mode), which reduces the height of the landing gear (and, therefore, reduces the relative weight of the chassis structure and glider) and lower- it has the aerodynamic drag of the aircraft (and, consequently, increases the aerodynamic quality of the aircraft); reduce the required wing area (since the wing area is determined by the conditions of balancing the aircraft in flight, and not by the requirements for the nose gear of the landing gear to take off), which increases the aerodynamic quality of the aircraft in cruise flight; to have a horizontal position of the air-propelled air intake on all flight modes of the aircraft, which allows one to have high efficiency of the air-propulsion wing (to have a high coefficient of recovery of the total pressure of the air-propulsion wing).

 By changing the sweep angle of the wing consoles of the claimed invention, in comparison with the prototype [12], it becomes possible to have a highly efficient take-off and landing mechanization on the wing (since the claimed invention has a center of pressure of the wing both in cruising and take-off and landing flight modes is in the same position with respect to the center of mass of the aircraft), which allows to radically reduce the wing area, which radically reduces the aerodynamic drag, and therefore, the radical Flax increases the maximum aerodynamic quality of the aircraft in cruising flight (both in comparison with the prototype and in comparison with other known aircraft of the aerodynamic schemes “tailless”, “flying wing” and “normal”).

The inventive aircraft has neither horizontal nor vertical plumage (that is, it is made according to the scheme “tailless”), which also increases its aerodynamic quality. The inventive aircraft has a straight, variable, sweep wing with a negative geometric twist and a negative transverse V angle. This allows us to provide static stability of the claimed aircraft in all axes only due to the wing: in pitch - due to sweep and negative geometric wing twists; along the roll - due to the required angle of the transverse V wing; at the rate - due to direct sweep, negative geometric twist and the required angle of the transverse V wing.

Brief Description of the Drawings

In FIG.1 4 shows one of the possible embodiments of the claimed invention in a variant of a subsonic administrative aircraft, where the numbers indicate: 1a, 16 and 1c - the center section of the wing in its position, with a horizontal cruise flight, on takeoff and landing flight modes and braking during landing, respectively; 2a, 26, and 2c show the right wing console in its position during horizontal cruise flight, in takeoff and landing flight modes and in braking mode during landing, respectively; For, 36 and Sv - the left wing console in its position, with horizontal cruise flight, on takeoff and landing flight modes and on the braking mode during landing, respectively; 4 - general engine nacelle; 5 and 6 - upper and lower double-circuit turbojet engines (turbofan engines), respectively; 7 - fuselage; 8 + 13 - wing elevons; 14 and 15 - fissile flaps of the wing; 16 + 21 - spoilers (dampers of the lifting force of the wing) of the wing; 22 and 23 - air intakes of the upper turbofan 5 and lower turbofan 6, respectively; 24 - dividing cheek air fighters 22 and 23; 25 and 26 fairings; 27 - bracket of a wing hinge; 28 - hydraulic cylinder for changing the wing center-wing installation angle; 29 - the front door of the fuselage 7; 30 - the upper influx of the fuselage 7; 31 - slot with a wedge for draining the boundary layer. Arrow with the inscription N.P. The flight direction of the aircraft is shown. The solid line shows the position of the center wing of wing 1a, right 2a, and left. For the wing consoles in horizontal cruise flight, the dash-dot line shows the position of the wing center section 16, the right 26 and left 36 wing consoles for takeoff and landing flight modes, and the dash-dot line shows the position of the wing center section 1c, the right 2c and the left Sv of the wing consoles during braking during landing.

 FIG. 1 shows a side view, FIG. 2 shows a front view, and FIG. H shows a top view of the inventive aircraft.

 Figure 4 shows the location of the center wing of the wing to the engine nacelle when viewed from the side.

 FIG. 5 shows a side view of one of the possible embodiments of the claimed invention in a variant of a subsonic administrative aircraft, where the numbers indicate: 35- fuselage; 36 - left wing console (the right wing console is not shown and not shown in the figure); 37 - wing center section; 38 and 39 - upper and lower turbofan engines, respectively; 40 and 41 - air intakes of engines 38 and 39, respectively; 42 and 43 are rotary all-angle nozzles of the engines 38 and 39, respectively.

FIG. 6 shows a front view of one of the possible embodiments of the claimed invention in a variant of a transport aircraft, where the numbers indicate: 48 — fuselage; 49 and 50 — left and right wing consoles, respectively; 51 - wing center section; 52, 53 and 54 - the left, central and right turbofan engines, respectively; 55 - common nacelle engines; 56, 57, 58 and 59 are vertical pylons.

Embodiments of the invention

The claimed invention, in one possible embodiment, in a variant of a subsonic administrative aircraft, is as follows. There is a swept (direct sweep) wing (FIG. 1 + 4) of variable sweep, consisting of a center wing (positions 1a, 16 and 1c), right (positions 2a, 26 and 2c) and left (positions Za, 36 and Sv ) consoles. There are two turbofan engines (upper 5 and lower 6), located in the common engine nacelle 4 one above the other in the plane of symmetry of the aircraft, offset from each other by some distance along the longitudinal axis of the aircraft (but may not be offset) , fuselage 7 (essentially a gondola for a payload). The wing center section (positions 1a, 16 and 1c) is hinged (the axis of the above hinge is perpendicular to the plane of symmetry of the aircraft) attached to the engine nacelle 4 (engine nacelle 4 is located on the underside of the wing center section). Thus, the wing center section (positions 1a, 16 and 1c) is made with the possibility of changing the angle of its installation (in the longitudinal plane) with respect to the nacelle 4 and the fuselage 7. The attachment center of the wing center and the rotary mechanism for changing the angle of its installation is made , for example, as in the case of well-known passenger aircraft, the attachment unit and the rotary mechanism for changing the angle of installation of the adjustable stabilizer (horizontal tail) are made, or is made as the attachment unit and the rotary mechanism for changing the angle of installation of the wing of the known vertical take-off and landing aircraft (VTOL) with a rotary wing, for example, as in the above American ex- 1 ">

of the XC-142A vertical VTOL (but there may be other acceptable designs of the assembly and the rotary mechanism - this is not a matter of principle, but of principle is that the wing center section has the ability to change its installation angle). The right (positions 2a, 26 and 2c) and left (positions Za, 36 and Sv) wing consoles are pivotally attached to the wing center section. At the same time, the right and left wing consoles have the ability to change their sweep angle within certain limits, turning in their above hinges. The attachment site of the rotary right and left wing consoles to the wing center wing (and the rotary mechanism for changing the sweep of the wing consoles) is made, for example, as in the known aircraft with variable sweep of the wing, for example, as in the aforementioned aircraft F-1 1 1 (but there may be other acceptable constructions of the assembly and the rotary mechanism - this is not important, but only the fact that the right and left wing consoles has a variable sweep). The attachment point of the left and right wing consoles to the wing center section and the rotary mechanism are not shown in the figures, since their specific design is not critical. The fuselage 7 is attached to the wing center section by means of a common engine nacelle 4 (ie, the fuselage 7 is attached to the underside of the engine nacelle 4, moreover, the fuselage 7 and the engine nacelle 4 are fixedly connected to each other).

Thus, the center wing (and, therefore, the entire wing, including the right and left wing consoles) is made with the possibility of changing the installation angle (in the longitudinal plane) with respect to the nacelle 4 and the fuselage 7. On the right and left consoles the wings have elevons 8 - ^ - 13, fissile flaps 14 and 15 (located at the ends of the wing consoles) and spoilers (dampers of the wing lift) 16-K21 (any other devices, for example, interceptors, can be used as wing dampers).

The wing in the claimed invention has a negative geometric twist and a negative angle to the transverse V. At the front of the nacelle 4 there are air intakes 22 and 23 of the upper 5 and lower 6 engines, respectively. The air inlet channels 22 and 23 are separated by a horizontal partition and have a separating cheek 24 (serves to exclude the influence of air flow in one air intake (for example, when surging in it) on the air flow in another air intake). Thus, the air intakes 22 and 23 of the engines 5 and 6 are located, on the one hand, on the lower side of the wing center, on the other hand, on the upper side of the fuselage 7 (that is, between the wing center section and the fuselage 7). Between the wing center section and the upper part of the engine nacelle 4 there is a gap for draining the side layer. In this drain slot there are two fairings 25 and 26, in which two wing link brackets 27 are placed (one bracket in each fairing) and two hydraulic cylinders 28, which serve to change the wing center angle in the longitudinal plane (one hydraulic cylinder in each fairing). Any other acceptable mechanism (one or several), for example, screw jacks, can be used as a device for changing the angle of the wing center wing in the longitudinal plane, for example, it’s not important, but only that - There is such a mechanism. The fuselage 7 has an entrance door 29 on the left side (on the right side of the fuselage 7 there is the same entrance door - not shown in the figures). On the upper side of the fuselage 7 (in the plane of symmetry of the aircraft) there is an influx of 30, increasing the height of the passenger cabin above the longitudinal passage. Between the influx of 30 and the the lower part of the engine nacelle 4 has a slot 31 with a wedge for draining the boundary layer.

 The claimed aircraft has neither horizontal plumage, nor vertical plumage - it is made according to the scheme tailless tailless.

The remaining units of the claimed aircraft do not affect the obtained technical result, and therefore are not indicated in the figures and in the description.

In a horizontal cruising flight (at subsonic speed), the wing center wing of the aircraft being installed is set, by means of two hydraulic cylinders 28, to the minimum installation angle (for example, equal to 3 °) with respect to the longitudinal axes of the engine nacelle 4 and the fuselage 7, and occupies the position 1a (shown in the figures by a solid line). The right and left wing consoles are set (by means of a rotary mechanism, which is not shown in the figures) to the maximum sweep angle (equal in magnitude to the right and left wing consoles), for example, equal to 30 ° (along the front edge of the wing) - as in well-known subsonic jet administrative planes, and occupy positions 2a and 3a, respectively (the figures show a dash-dot line). In cruising flight, the right 2a and left Behind the wing console have a direct sweep. At the same time, the fuselage 7 and the engine nacelle 4 maintain their horizontal position, and the wing center section (and, consequently, the entire wing) is set to the optimal angle of attack. The center of pressure of the wing (the point of application of the lifting force of the wing) is in such a position with respect to the center of mass of the aircraft, which provides the required (optimal) degree of static stability of the aircraft in pitch.

When landing (and during take-off), the wing center section of the claimed aircraft is installed, by means of two hydraulic cylinders 28, on the take-off the landing angle of the installation in the longitudinal plane (for example, at an angle of 10 °) with respect to the longitudinal axes of the engine nacelle 4 and the fuselage 7, and occupies position 16 (the figures show a dash-dot line). That is, the wing center section rotates about an angle of 10 ° relative to the axis of its hinge (located in the brackets of the wing 27). At the same time, the fuselage 7 and the engine nacelle 4 maintain their horizontal position. The right and left wing consoles are installed (by means of a rotary mechanism, which is not shown in the figures) at the minimum sweep angle (the same in magnitude for the right and left wing consoles), and occupy positions 26 and 36 (the dashed-dotted diagram in the figures line), respectively. When landing (and taking off) the claimed aircraft, its right 26 and left 36 wing consoles have a direct sweep. Elevons 8-43 of the wing bow to positive angles of attack, thus fulfilling the function of take-off and landing mechanization (increasing the coefficient of lift of the wing). When landing (and taking off) the wing consoles 26 and 36, turning relative to the axes of their hinges, reduce their sweep angle to the desired value, which is determined by the fact that the center of pressure of the wing (with the wing operating and operating mechanics) is in the same position with respect to the center of mass of the aircraft as in cruising flight, thus providing the desired (optimal) degree of static pitch stability of the aircraft.

 Thus, in the claimed invention, the center of pressure of the wing both in cruising flight and in take-off and landing flight modes is in the same position with respect to the center of mass of the aircraft (but can be located in different places — if necessary).

Reduced wing sweep on takeoff and landing modes flight (compared with the angle of sweep of the wing in cruising flight) allows the claimed aircraft (without horizontal tail) to have highly efficient take-off and landing mechanization (of any acceptable type, and not just its variant considered above). This allows the claimed invention to radically reduce the wing area, therefore, radically reduce the aerodynamic drag, and, therefore, radically increase the maximum aerodynamic quality of the aircraft, including in cruise flight (compared with known aircraft without horizontal feathering - aerodynamic schemes “tailless” and “flying wing”). The wing of the claimed invention may have a maximum lift coefficient during take-off and landing of the same magnitude as that of aircraft of the “normal” aerodynamic design. Thus, the claimed invention allows to realize all the potential advantages of an aircraft without horizontal tail (aerodynamic tailless design) in terms of maximizing the aerodynamic quality of the aircraft.

As can be seen from the foregoing, the sweep of the wing in the claimed invention changes for a completely different purpose (compared with the known aircraft with variable sweep of the wing, for example, compared with the above aircraft F-1 1 1). Namely, in order for the wing pressure center both in cruise flight and in take-off and landing flight modes (when the take-off and landing wing mechanization, which increases the wing lift coefficient), be in the same position, with respect to to the center of mass of the aircraft. In this case, the sweep of the wing consoles of the claimed invention will need to be changed by an angle of about 10 °. This is because when changing the sweep of the wing. the center of mass of the aircraft moves (in the direction of the longitudinal axis of the plane) by a small amount (about 2% of the average aerodynamic chord (SAX)), while the center of pressure of the wing moves (in the direction of the longitudinal axis of the plane) by much larger value (when changing the sweep angle of the wing consoles by about 10 ° - it moves by 25% of the MAR).

 And for known aircraft with variable wing sweep and for the claimed invention, the wing bearing capacity (wing lift coefficient) increases during take-off and landing flight modes due to a change (decrease) in the wing sweep angle. But this happens for completely different reasons. In well-known aircraft with variable wing sweep (which were all supersonic airplanes of the “normal” aerodynamic design, in which the arrow sweep of the wing consoles changed by tens of degrees - from 15 ° to 72 °), the wing lift coefficient increases only by reducing the angle of sweep of the consoles wings. In the claimed invention, the wing lift coefficient increases not so much due to a decrease in the sweep angle of the wing, but due to the fact that it becomes possible for a tailless aircraft to have, in principle, a highly efficient take-off and landing mechanism.

The inventive aircraft is controlled by: pitch and roll - by deflecting the elevons 8-43; along the course - by deflecting the fissile flaps 14 and 15 located at the ends of the right and left wing consoles (for example, as is the case with the famous American B-2 bomber made according to the flying wing aerodynamic scheme). At the same time, it is possible to control the claimed aircraft by pitch when the front elezons 12 and 13, on the one hand Nines, and the rear elevons 8 and 9, on the other hand, deviate differentially (deviate in different directions).

 The use in the claimed invention of a wing with a variable installation angle (in the longitudinal plane, relative to the fuselage and the engine nacelle) allows, in comparison with the prototype [12]: to maintain the horizontal position of the fuselage and engine nacelle (in any flight mode), which reduces the height of the landing gear (and, therefore, reduces the relative weight of the chassis and airframe) and reduces the aerodynamic drag of the airplane (and, therefore, increases the aerodynamic quality of the airplane); reduce the required wing area (since the wing area is determined by the conditions of balancing in flight, and not by the requirements for the landing gear nose wheel to take off), which increases the aerodynamic quality of the aircraft; The horizontal position of the engine nacelle in all flight modes makes it possible to have high engine efficiency (to have a high value of the coefficient of restoration of full engine pressure). The location in the claimed invention of the engine air intakes under the wing allows for high engine efficiency (high value of the recovery coefficient of the total pressure of the engines) when changing (increasing) the wing center angle.

The use of a wing with a variable sweep angle in the claimed invention allows, in comparison with the prototype [12] (and compared with other well-known aircraft aerodynamic schemes "tailless", "flying wing" and "normal"): to have a highly efficient wing take-off and landing mechanization (as one of the possible options, elevons deflected by positive angles - as in the variant considered above), which allows to drastically reduce the required wing area, which drastically reduces the aerodynamic drag of the wing, and, consequently, drastically increases the maximum aerodynamic quality of the aircraft during cruise flight. At the same time, the claimed aircraft can have wing take-off and landing gear of any acceptable type (flaps (single and multiple-winged); flaps; slats; energy mechanization (suction or blow-off of the boundary layer); combination of the above; and others).

 Thus, the claimed invention allows to realize all the potential advantages of the aircraft of the aerodynamic scheme "tailless" (and the aircraft as such) in terms of maximizing its aerodynamic quality.

 The inventive aircraft between the lower surface of the wing center section and the upper surface of the engine nacelle 4 has a slot for draining the boundary layer. Two fairings 25 and 26 are installed in the slit. In this case, each fairing consists of two parts (not shown in the figures) - the upper part is fixedly attached to the lower surface of the wing center section, and the lower part is fixedly attached to the upper surface of the engine nacelle 4. When changing the wing center angle, the upper and lower parts of the fairings 25 and 26 move telescopically relative to each other, thereby preserving the streamlined shape of the fairings 25 and 26. When changing the wing center angle, the shape of the bow ntroplana wing is not changed, that-ground for aerodynamic efficiency claimed aircraft,

The claimed invention has two engines (turbofan engines) located one above the other in the plane of symmetry of the aircraft at a minimum distance from each other. Consequently, if one of the engines fails, there is no destabilizing moment along the course, which makes it possible to abandon the vertical plumage as such. It reduces aerodynamic drag and increases the aerodynamic quality of the aircraft as a whole. The necessary static stability of the declared aircraft along the course is ensured by the direct sweep of the wing and the angle of the transverse V wing.

 In such an embodiment of the claimed invention, the function of the pylon by which the fuselage is attached to the wing is performed by the engine nacelle. At the same time, the construction height of the nacelle (in the horizontal plane) is quite sufficient to eliminate the problems associated with the aeroelasticity of the airframe structure.

 An embodiment of the claimed invention is possible when it has a thrust vector of the engine (or engines) can change its direction, for example, in the longitudinal plane relative to the wing chord (to create a pitch moment - for balancing and controlling the aircraft). This can be done, for example, by using a rotary nozzle on the engine.

 In the claimed invention, as a device that creates traction force (i.e., as a propulsor), any of their acceptable types can be used: turbofan engines of any type (single-circuit turbojet engines (turbojet engines), double-circuit turbojet engines, ramjet engines) ), and others); liquid propellant rocket engines (LRE); a propeller or fan (pulling or pushing) driven by an engine (or motors) of any acceptable type (piston, turboshaft, electric, and others), while the fan (and propeller) can be as open Gym, and placed in the channel; and other.

An embodiment of the claimed invention is possible when each of the two engines is combined, for example, a combination of a turbojet engine and ramjet engine. An embodiment of the claimed invention is possible when one of the engines (for example, the upper one) is a turbojet engine and the other (lower) is a ramjet engine.

 An embodiment of the claimed invention is possible when, in addition to (or instead of) the engine, another type of engine is used, for example, an engine.

 The invention may have one or more engines located in the plane of symmetry of the aircraft one above the other (which may either be offset from each other in the direction of the longitudinal axis of the aircraft, or not offset).

 A possible embodiment of the claimed invention with a different (compared to that shown in FIG. K4) arrangement of engines (engines), for example, two engines are located on horizontal pylons in the rear of the fuselage (as in the well-known subsonic jet administrative aircraft) and the wing center wing is hinged directly to the fuselage, all other things being equal.

A possible embodiment of the claimed invention is different from that shown in FIG. H4 in that instead of two turbofan engines 5 and 6, two fans are installed, driven, for example, by electric motors, ceteris paribus.

 In the claimed invention, a device (or several devices) generating a traction force can be located in any suitable place on the aircraft.

An embodiment of the claimed invention is possible when it generally does not have a device generating traction force (propulsion), while its wing center section is pivotally attached directly to the fuselage (for example, from its upper side), ceteris paribus. That is, in this embodiment, the claimed invention is made in the form of a glider.

 In the claimed invention, the wing can have any acceptable shape in terms of: direct or reverse sweep; small elongation; high elongation; and other. For example, an embodiment of the claimed invention is possible when the wing has a direct sweep in its cruise flight, and the wing has a reverse sweep during take-off and landing flight modes.

 The use of a variable wing sweep and a variable thrust vector of the engines of the claimed invention makes it possible to perform it in a variant of a short take-off and landing airplane, in which at take-off and landing the engines create a vertical component of the thrust force.

In the claimed invention, the change in the sweep of the wing can also be used to expand the range of centerings of the aircraft - which eliminates one of the disadvantages of the known aircraft tailless aerodynamic design.

 The inventive aircraft can be performed according to any acceptable aerodynamic scheme: tailless (as discussed above), duck, normal, and more.

 A possible embodiment of the claimed invention differs from that shown in FIG. 1-M in that it has a vertical tail of any suitable type, for example, in the form of end washers (or wings) mounted on the ends of the wing consoles.

 The claimed invention can be used: as a manned aircraft of any type (for example, in the embodiment of a passenger aircraft); as an unmanned aerial vehicle.

The inventive aircraft can have any acceptable flight speed: subsonic; supersonic; hypersonic. 00934

23

 In a variant of a supersonic (or hypersonic) aircraft of the claimed invention, in a supersonic engine air intake as a generator of shock waves, a vertical wedge common to both engines can be used (as one of the possible options) (for example, a multi-stage, unchanged geometry) . The air intake channels of the engines are separated by a horizontal partition. In this embodiment of the claimed invention, at supersonic flight speeds, the shock waves from the vertical wedge and the air intake shell of the engine nacelle sit only on the lower surface of the wing, which allows to reduce the angle of attack of the wing, and, therefore, to reduce the aerodynamic drag of the wing and increase the aerodynamic aircraft quality.

 In a variant of a supersonic or hypersonic aircraft of the claimed invention, when flying at a supersonic or hypersonic speed, the shock waves from the nose and side of the fuselage sit only on the lower surface of the wing, which increases the aerodynamic quality of the aircraft as a whole. That is, in the claimed invention there is a useful supersonic and hypersonic interference between parts of the aircraft.

 The claimed invention can be used both in a variant of a plane for ordinary take-off and landing, and in a variant of a plane for vertical take-off and landing (for example, with a vertical arrangement of the fuselage during take-off and landing).

An embodiment of the claimed invention is possible, which differs from that shown in FIGS. 1- ^ 4 in that its right and left wing consoles are capable of differentially changing their installation angles (for example, the right wing console increases its angle 00934

24

installation, and at this time, the left wing console reduces its installation angle (like some well-known aircraft, the consoles of the fully rotatable horizontal tail differentially deviate)).

 In the claimed invention, a change in the sweep of the wing can be used to reduce the size of the aircraft when it is stored in the hangar or in the hold of the ship.

 In the claimed invention, the mechanisms for changing the wing installation angle and changing the wing sweep can be of any acceptable type: hydraulic drive; electric drive; and other.

A possible embodiment of the claimed invention (FIG. H4), when his wing is used as a means of reducing the mean free path of the aircraft during landing. In this embodiment, when landing the claimed aircraft, after the wheels touch the landing gear, the wing is organized by deflecting the spoilers 16 - ^ - 21 (or in any other suitable way, for example, by means of spoilers) located on the consoles wing, decrease (damping) of the aerodynamic lift force of the wing and increase the force of aerodynamic drag of the wing. That is, a stall is organized on the wing. Then the wing center wing (and, consequently, the right and left wing consoles), by means of two hydraulic cylinders 18, are set to positions 1c, 2c and Sv (shown in figures as a dashed line), respectively, for example, at an angle of 90 ° (but they can be installed at any other acceptable angle), with respect to the longitudinal axes of the engine nacelle 4 and the fuselage 7. That is, the angle of the wing center section during airplane braking during landing is increased compared to the take-off and landing angle (when wing there is no stall). This leads to a further decrease in (extinguishing) the aerodynamic lifting force of the wing and increasing the force of the aerodynamic drag of the wing, which is used to reduce the path length of the claimed aircraft during landing. At the same time, the fuselage 7 and the engine nacelle 4 retain their horizontal positions.

 Thus, due to an increase in the installation angle (angle of attack) of the wing (in the longitudinal plane), compared with the take-off and landing angle of its installation, the flow stalls on the wing, as a result of which the force of the aerodynamic drag of the wing increases sharply, which reduces the length of the run when landing.

 The use of a wing in the claimed invention to reduce the flight path of an airplane during landing, in comparison with known airplanes with brake parachutes, has the following advantages: it creates a greater aerodynamic drag force (since the wing area on the aircraft is larger than the area used on aircraft brake parachutes), which reduces the flight path of the aircraft during landing and allows the claimed aircraft to land on airfields with a shorter runway length; the wing does not need to be dropped from the plane and quickly removed from the landing strip; the aircraft uses the mechanism that already exists on it to change the angle of the wing (you do not need to have an additional unit such as a brake parachute - but you can also have one).

In the claimed invention for using the wing as a means to reduce the path length during landing, the angle of installation of the wing may increase (in absolute value, compared with the take-off and landing angle of the installation), or in the direction of increasing the positive angle of the installation (as in the case considered above, when the wing creates, albeit not large (due to the stall of the flow on it), but PT / RU2015 / 000934

26

specific lifting force (but may not create)), or towards the negative installation angle (in this case, the wing creates a small negative lifting force (but may not create) and a large aerodynamic drag force).

An embodiment of the claimed invention (FIG. 5) is possible, for example, in an embodiment of an administrative aircraft, when it has two engines 38 and 39 located one above the other in the plane of symmetry of the aircraft in the rear of the fuselage 35. Center wing wing 37 hinged attached to the air intakes 40 and 41 of the engines 38 and 39, and made with the possibility of changing the installation angle (with respect to the fuselage 35). The left 36 and right (not shown in the figure) wing consoles are pivotally attached to the center section of the wing 37 and are configured to change their sweep angle (with respect to the center plan 37 and the longitudinal axis of the fuselage 35). The air intakes 40 and 41 of the engines 38 and 39, respectively, are located on the upper side of the fuselage 35 (and on the lower side of the wing center section 37), and are separated by a horizontal partition. The air intakes 40 and 41 are fixedly attached to the fuselage 35. All-angular rotary nozzles 42 and 43 of the engines 38 and 39, respectively, allow you to change the thrust vector of the engines in the vertical and horizontal planes, which is used to control the aircraft in pitch and course. At the same time, the use of the claimed invention of all-angular rotary nozzles 42 and 43 of the engines 38 and 39 allows to radically increase the flight safety of the claimed aircraft (compared to known aircraft), since even when the claimed aircraft (when the flow on the wing is interrupted) is in dangerous flight modes (stall, corkscrew), by changing the thrust vector of the engines, the claimed aircraft can be brought out of these dangerous flight modes. An embodiment of the claimed invention (FIG. 6) is possible, for example, in the embodiment of a transport aircraft when it has three engines 52, 53 and 54 located horizontally in a common engine nacelle 55 on the upper side of the central part of the fuselage 48. The wing center section 51 - is complete with the possibility of changing the installation angle (with respect to the fuselage 48), and the left 49 and right 50 of the wing console are pivotally attached to the wing center section 51 and are configured to change their sweep angle (with respect to the center section 51 and the longitudinal axis the fuselage 48). The nacelle 55 by means of two vertical pylons 56 and 57 is pivotally attached to the center section of the wing 51 from its lower side. The engine nacelle 55 is fixedly attached to the central part of the fuselage 48 from its upper side by means of two vertical pylons 58 and 59. Thus, the air intakes of the engines 52, 53 and 54 are located on the upper side of the central part of the fuselage 48 and on the lower side of the wing center section 51.

Industrial applicability

The claimed invention can be used as: an aircraft of any acceptable type (supersonic, hypersonic, subsonic), both manned (for example, passenger) and unmanned); glider.

Claims

CLAIM 1. Aircraft (LA) of the tailless aerodynamic scheme, has a fuselage, a wing, characterized in that the wing consists of a center section, left and right consoles, the center section of the wing is configured to change the angle of its installation (in longitudinal plane) with respect to the fuselage, the left and right wing consoles are pivotally attached to the wing center section and are configured to change the angle of their sweep, the wing has a take-off and landing mechanism.  2. LA popL, characterized in that it has at least one device made to create traction.  3. The aircraft according to claim 2, characterized in that an air-jet engine (WFD) placed in the plane of symmetry of the aircraft is used as a device for creating traction force, the air intake of the above WFD is located on the lower side of the wing center section and on the upper side fuselage, the above-mentioned wing is made direct sweep.  4. Aircraft according to claim 3, characterized in that the above WFD has a rotary nozzle. 5. Aircraft according to claim 3, characterized in that it has two WFDs located in a common engine nacelle one above the other in the plane of symmetry of the aircraft, the above WFDs have rotary nozzles, the air intakes of the above WFDs are located on the lower side of the price troplana wing and the upper side of the fuselage. 6. The aircraft under item 3, characterized in that. that has at least two WFDs located in a common engine nacelle in a horizontal plane, the air intakes of the above WFDs are located on the lower side of the wing center section and on the upper side of the fuselage.  7. The aircraft according to claim 2, characterized in that at least one fan is used as a device made to create traction force, placed in the plane of symmetry of the aircraft in the air duct and driven, for example, by an internal combustion engine or an electric motor, the entrance to the aforementioned air channel is located on the lower side of the wing center section and on the upper side of the fuselage, the above wing is made direct sweep.  8. LA according to claim 2, characterized in that at least one propeller located in the plane of symmetry of the aircraft and driven, for example, by an internal engine, is used as a device made to create traction force combustion or electric motor, the above wing is made direct sweep.  9. Aircraft according to any one of claims 1 to 8, characterized in that it has a means to reduce the path length during landing, the quality of which the above wing was used by installing it on a large, in absolute value, installation angle, than the landing angle of its installation.