RU2384472C2 - Airplane, fast-speed arrow wing and fuel system - Google Patents

Abstract

FIELD: air transport.

SUBSTANCE: airplane includes nose, front, central, rear and tail sections and fast-speed arrow wing with jet engines. Fast-speed arrow wing is characterised by change of maximum relative profile thickness, geometric twist angle, angle of lateral V wing under wing semispan and middle surface curvature both along semispan and profile chord. Fuel system includes reservoir of the airplane central section and reservoirs in outer wings near the central section at middle part and at the periphery. The central section is equipped with expandable walls. Fuel system provides initial use of fuel from central section reservoir, then from fuel tanks of outer wings located near central section and then from tanks at the periphery of outer wings, and after all the others - from the middle part of the wing by fuel transfer between tanks.

EFFECT: improvement of aerodynamic characteristics for planes with different weight and hull length.

22 cl, 54 dwg

Description

The claimed group of inventions relates to aviation technology, namely to passenger aircraft, and includes technical solutions united by a single inventive concept, allowing to solve the problems of choosing design solutions for a passenger aircraft designed to develop on its basis a family of passenger aircraft - a model range of aircraft with different passenger capacities and the maximum estimated range, as well as the aerodynamic design of its high-speed swept wing and its fuel with tem.

Analysis of technical solutions related to the choice of the appearance of a passenger aircraft to create a family of passenger aircraft on its basis, and the dynamics of their development (see, for example, Designing passenger aircraft, edited by G.V. Novozhilov, M .: Engineering, 1991 , pp. 244-249) allows us to distinguish two approaches to the development of a model range of passenger aircraft.

The first approach is associated with the sequential creation of a family of aircraft of the family for a long time, when each subsequent aircraft model is created on the basis of the previous one. Examples of the implementation of this approach to creating a family of passenger aircraft is the process of developing families of passenger aircraft Il-18, Il-62, Boeing-707, Boeing-727, Boeing-737. The second approach is associated with the initial design of an aircraft designed to develop a family of aircraft on its basis, that is, with the parallel design of a family of aircraft, including aircraft of different passenger capacity and / or purpose with the maximum degree of unification. Examples of the implementation of this approach can be found in the decisions of the patent of the Russian Federation 2148534, Germany 4323020, USA 6575406, 6938854, EPO 0857648, in which the family aircraft differ in the shape of the fuselage, the number of engines and their characteristics, the number of wings. The practical implementation of these technical solutions is quite problematic.

Examples of practical implementation of the second approach to the selection of a family of aircraft of the family include the family of passenger aircraft L-1011, DC 10, А319-А321, which are based on the design of an aircraft that serves as the basis for the entire model range of aircraft of the family.

Aircraft L-1011 (see PN Belyanin. Production of wide-body aircraft. M: Mechanical Engineering, 1979, pp. 87-101) contains a fuselage, high-speed swept wing, vertical and horizontal tail, engines and chassis.

The L-1011 airplane is divided by technological joints into the nose of the fuselage (nose compartment), the middle of the fuselage (front compartment, central compartment), the rear compartment, the tail compartment, the detachable wing consoles, the center section with additional walls parallel to the side members, vertical and horizontal tail . The high-speed swept wing includes the front and rear spars and a power set of stringers and ribs.

The model range of the Lockheed Aircraft L-1011 family aircraft (see BV Belyaev, Passenger Aircraft of the World. M .: Aspol, Argus, 1997, pp. 236-242) includes two groups of aircraft, the first of which It is designed for transportation of up to 400 passengers, the second for transportation of up to 330 passengers.

The first group of aircraft contains aircraft designed for four levels of flight range. The plane with the lowest level of range of this group is the L-1011-1 aircraft: when transporting 276 passengers with baggage, its flight range is 5350 km. Aircraft L-1011 - 100 and L-1011-200 are models of aircraft with intermediate levels of flight range, with a similar load of the flight range of these aircraft are 6390 and 7600 km. Aircraft L-1011-250 is the aircraft with the highest level of flight range in this group, its flight range for carrying 276 passengers with baggage is 8890 km.

The second largest passenger group of aircraft of this family is represented by one model - L-1011-500, designed to carry 246 passengers with baggage with a flight range of 9700 km.

The transition from the first group of aircraft of the family to the second is carried out by reducing the length of the front and rear compartments, in addition, when moving from the aircraft of the first group to the second, the wing span increases.

In the internal volume of the high-speed swept wing, divided by technological joints into the center section and detachable wing consoles, fuel tanks are placed. In this case, the transition within the first group of aircraft from an aircraft with the shortest flight range L-1011 - 1 to an aircraft with an intermediate flight range L-1011 - 100 is carried out by introducing a central fuel tank into the center section design and a corresponding increase in fuel mass. The transition to the next level of flight range (aircraft L-1011-200) from the aircraft L-1011-100 was carried out through the use of more powerful engines, and the transition to the aircraft with the longest flight range L-1011-250 was carried out due to a further increase in the fuel mass in central fuel tank.

Another example of a technical solution for an aircraft designed to develop a family of passenger aircraft is the DC-10 aircraft of the company Donnel Douglas.

The aircraft of the family in question (see PN Belyanin. Production of wide-body aircraft, Moscow: Mashinostroenie, 1979, pp. 72-87) consists of a fuselage divided by technological joints into the bow, front, center, rear and tail compartments, high-speed swept wing, chassis.

The high-speed swept wing consists of a center wing and detachable wing consoles, made according to the two-spar design, and includes the front and rear side members, the skin with its upper and lower parts, ribs and a power set of stringers. The wing is made with a wing half-span variable relative thickness of the wing profiles, while the relative thickness of the wing in the root part is ~ 12.2%, in the end part ~ 8.4%. The sweep of the wing along the line of the quarters of the chords is 35 degrees.

Under the wing consoles are two engines in the nacelles. In addition, it is possible to place another engine in the rear of the fuselage. In addition, engines of various brands with different characteristics can be used in an aircraft.

The fuel system includes wing fuel tanks and a center section fuel tank. On the basis of this aircraft, a family of aircraft can be developed, including 2 groups of aircraft with different passenger capacities.

The first group, designed to carry 236 passengers, includes one aircraft (DC-10 Twin), designed for flights up to 4050 km (see Strukov Yu.P. Modern Aircraft of the USA and Western Europe, Results of Science and Technology, Ser. Aircraft, t .1, M., VINITI, 1973, p. 47).

The model range of the second largest passenger group of this family, designed to carry up to 380 passengers, includes 4 models of aircraft with different levels of flight range (see Belyaev V.V. Passenger Airplanes of the World. M: publishing house Aspol, Argus, 1997. , pp. 262-269). In this group, the DC-1010 has the shortest flight range of 5370 km when transporting 277 passengers with baggage, the DC-1015 is a model with an intermediate level of range of this group - 6940 km when transporting the same number of passengers. The DC-1030 and DC-1040 aircraft are the models with the longest flight range in this group: when transporting 277 passengers with baggage, the flight range of the DC-1030 is 10,460 km, and the DC-1040 is 10,280 km.

In the family of aircraft under consideration, the transition from one group in terms of passenger capacity to another group is carried out by changing the fuselage length and changing the number of engines on the aircraft: the fuselage length of the first group in terms of passenger capacity is less than the length of the fuselage of the second group in terms of passenger capacity, while the aircraft of the first group in terms of passenger capacity is equipped two engines, and the second in terms of passenger capacity of the group - three.

Within the group's second largest passenger capacity, the transition from an aircraft with the smallest estimated maximum range (DC-1010) to aircraft with an intermediate (DC-1015) and the largest estimated maximum range (DC-1030, DC-1040) was made due to the following design changes aircraft with the smallest estimated maximum range:

a change in the wing of the aircraft (the wingspan of the aircraft is increased by 3.05 m);

a change in the fuel system of the aircraft (the volume of fuel tanks was increased due to the introduction of a central fuel tank in the center section);

a change in the chassis of the aircraft (aircraft with an intermediate and maximum estimated range are equipped with an additional (third) main landing gear located under the fuselage between the two main pillars);

a change in aircraft engines with different traction characteristics and fuel economy indicators.

The above technical solutions for aircraft are suitable for the development of families of wide-body aircraft, designed to carry from 200 to 400 passengers with a range of flight ranges, including intercontinental flights. In addition, the model range of the considered aircraft families, represented by two groups of aircraft with different passenger capacities, one of which includes only one aircraft model with one level of maximum estimated flight range, cannot in modern conditions provide a flexible response to changing airline requirements for passenger aircraft with different passenger capacity and flight range. This is especially true for the segment of regional passenger aircraft, designed to carry from 60 to 100 passengers.

The considered technical solutions for the model series of aircraft families are characterized by a low degree of unification of aircraft: aircraft can use various engines, for example engines of various manufacturers, the design of large units of the airframe (wings, landing gear, fuel tank structure, etc.) can be changed.

The solution to the problem of designing a family of aircraft with a high degree of unification, including groups with different passenger capacities, and therefore various longitudinal dimensions of the fuselage, requires solving the problem of equipping family aircraft with a fast swept wing of the same geometry.

A technical solution is known for a high-speed swept wing (see RF patent 1775972, IPC ВСС 3/10, publ. 09/30/1994). In accordance with this technical solution, each of the wing consoles is composed of two parts delimited from each other by basic sections parallel to the plane of symmetry of the aircraft, placed sequentially along the span of the wing consoles, seamlessly mated with each other and formed as a single spatial system. In this solution, the first part of each of the wing consoles is limited by the side section and the section in the middle part of the wing console at z = 0.3, the second part of each of the wing consoles is limited by the marked section in the middle part of the wing console and the end section.

The wing consoles are designed as a single spatial system based on a non-planar median surface with a relative profile thickness decreasing along the wing half-span and decreasing during the transition from the side to the end section, the angle of the geometric twist of the wing sections.

When passing from the side to the wing end section, the section profiles are made with a monotonic decrease in the maximum relative profile thickness from 16 ... 17% in the side section to 15.6 ... 16% at the boundary of the first and second parts and up to 13% in the end section.

до

и далее при переходе к концевому сечению до

.In addition, in this technical solution, the wing is made with a negative angle of the geometric twist of the sections, decreasing in wing span: when changing from the side section to the boundary between the first and second parts of the wing consoles, the angle of installation of the sections changes from

before

and then on going to the end section to

.

, а при переходе от бортового сечения к границе между первой и второй частями консолей крыла максимальная относительная вогнутость профилей возрастает от

до

и при дальнейшем переходе по второй части аэродинамической несущей поверхности практически не меняется.In addition, in this technical solution, the profiles of the first and second parts of the wing consoles are made with the location of the midline of the profiles above the chord. In this technical solution, as well as in the technical solution considered below for the patent of the Russian Federation 2228282, the position of the midline of the profile relative to the chord is considered as one of the signs. Hereinafter, if the midline of the profile is located above the chord, we will use the term positive concavity of the profile, and when designating the maximum relative positive concavity (the ratio of the maximum deviation of the middle line from the chord to the length of the chord), use positive values. If the midline of the profile is located below the chord, we will use the term negative concavity, and negative values will be used to indicate the maximum negative concavity. The maximum relative positive concavity of the profiles on the first and second parts of the consoles in this decision of the swept wing is placed at

, and when passing from the side section to the boundary between the first and second parts of the wing consoles, the maximum relative concavity of the profiles increases from

before

and with further passage along the second part of the aerodynamic bearing surface, it practically does not change.

, постоянной стреловидностью по передней и задней кромкам со стреловидностью по 1/4 ј хорд χ_1/4=3°, при этом стреловидность по линии заднего лонжерона χ_зл=0°.In accordance with this technical solution, a high-speed swept wing is used for the wing of a subsonic aircraft with lengthening λ = 11 ... 12, narrowing

, constant sweep along the front and rear edges with sweep along 1/4 ј chords χ _1/4 = 3 °, while sweep along the line of the rear spar χ _PL = 0 °.

This technical solution of the high-speed swept wing is effective at cruising flight speeds with a Mach number of up to 0.64 ... 0.66 and does not provide optimal aerodynamic characteristics when flying at a cruising speed of up to 0.82 M. In addition, this technical solution does not solve the problem of placement under wing consoles of turbojet engines of a large degree of bypass, characterized by large dimensions.

The development of the aircraft to create on its basis a unified model range of the family of passenger aircraft poses certain problems in the development of the fuel system of the family aircraft. At the same time, to the traditional technical problems that are solved during the development of aircraft fuel systems, for example, increasing the mass of refueling fuel, reducing the dry weight of fuel systems, decreasing undeveloped fuel residues, optimizing fuel production from multi-tank fuel systems of airplanes, adding tasks to develop a fuel system optimized for use in aircraft designed for different passenger capacities and different maximum estimated flight ranges.

An example of a well-known technical solution to the fuel system of an airplane is a technical solution to the fuel system of an airplane TU-134A (see V.A. Borodenko. “Aircraft TU-134A”, publishing house “Mashinostroyenie”, 1975, pp. 44-51, 305- 308, 319). This fuel system contains fuel tanks located in the center section and wing consoles. In this technical solution, the wing contains a longitudinal power set, including the front and rear spars, a transverse power set with ribs and a casing with its upper and lower parts. The center wing, made in the form of a caisson, is bounded by the front and rear spars, side ribs, upper and lower panels of the skin and is divided by ribs into four compartments. Each of the wing consoles is made in the form of a caisson, bounded by the front and rear spars, lower and upper skins and is equipped with ribs, some of which are sealed, and a longitudinal power set. The wing in this technical solution is made with a negative angle of the transverse V wing.

The fuel system of this technical solution contains two groups of fuel tanks located in the right and left consoles of the wing of the aircraft, each of which includes the first fuel tank (in tank terminology - tank No. 1) adjacent to the center section of the aircraft, and the second fuel tank (in source terminology - tank No. 3), located on the periphery of the wingspan, and a third fuel tank (in the terminology of the source, tank No. 2), located between the first and second tanks, and a consumable compartment, located in the internal volume of one of the fuel tanks, giving fuel to the engine, the input of which is connected to the consumable compartment. In addition, the fuel system is equipped with fuel transfer lines with transfer pumps, the inputs of which are connected to the internal cavities of the wing tanks, and the outputs - to the consumable compartments of the groups of tanks.

In this technical solution of the fuel system, consumable compartments are located in the internal volumes of the first fuel tanks, while one of the walls of the consumable compartments is combined with the front spar, and the other two with ribs located inside the first tanks of the fuel system. In addition, two additional small fuel tanks are placed in the two central sections of the center section (in the terminology of the source of tanks No. 1A), one of which is structurally included in one and the other in the other group of fuel tanks.

When developing fuel in flight, first the second fuel tanks located on the periphery of the wing consoles are emptied, then about half of the fuel from the first fuel tanks adjacent to the aircraft fuselage, then the third fuel tanks, after which the second half of the fuel from the first fuel tanks is generated and last fuel is produced from the consumable compartments. To control the processes of fuel production, the system provides for the presence of complex fuel production control systems.

The fuel system of this technical solution does not ensure its effective use in an aircraft intended for use in the model range of aircraft with different passenger capacity and different maximum estimated flight ranges, differing in different values of the maximum fueling. The mass of the fuel system is also great; during the flight, the position of the center of gravity of the aircraft changes significantly.

The closest analogue to the claimed technical solution "Airplane" is the technical solution of the aircraft, which is used in the family of passenger aircraft A319-A320-A321.

An airplane of this family (see Passenger aircraft Airbus Industry A320, compiled by Zaitseva N.N., TsAGI, 1993, pp. 13-48, pp. 51-58) is equipped with a fuselage divided by technological joints into successively placed forward, front , central, rear and tail compartments. The nose, center and tail compartments of the fuselage of the aircraft are identical for all aircraft of the family, and the front and rear compartments are made with a length that makes it possible to place on board the number of passengers corresponding to the group of aircraft in terms of passenger capacity. Due to this, the length of the aircraft changes and the transition from one group in terms of passenger capacity to another is carried out: for example, the length of the aircraft of the first group is 33.84 m, the second - 37.57 m, the third - 44.51 m.

относительная толщина профилей уменьшается, достигая локального минимума при

, а от

до

увеличивается, после чего при

относительная толщина профилей продолжает уменьшаться.All models of aircraft of this family are equipped with an identical for all aircraft of the family high-speed swept wing of the same geometry and one design. In this case, the wing is made with a sweep angle of 25 °, elongation λ = 9.42 with a constant positive angle of the transverse V wing, close to 5 ° 6'36 ''. When passing from the side to the end section, the relative thickness of the profile decreases from 15.3% to 10.8%. At the same time, on the half-span of the wing console adjacent to the side rib, the wing console is equipped with a section with a local minimum of the relative thickness of the profiles (see Passenger aircraft Airbus Industry A320. TsAGI, Moscow, 1993, p. 22, Fig. 26): from the side rib to the relative half-span coordinate

the relative thickness of the profiles decreases, reaching a local minimum at

, and from

before

increases, after which at

the relative thickness of the profiles continues to decrease.

The high-speed swept wing of family aircraft consists of a center section and wing consoles. It includes a longitudinal and transverse power set, while the front and rear spars and stringers are included in the longitudinal power set, and the ribs, some of which are sealed, and the wing skin with its upper and lower parts are included in the longitudinal power set.

In each of the wing consoles, a dry compartment is placed, while one of the walls of the dry compartments is aligned with the front spar. In the considered technical solution, the other two walls of the dry compartment are aligned with the ribs of the wing, and the fourth wall separates the dry compartment from the internal volume of the wing console.

In addition, the aircraft of the family are equipped with a power plant containing two turbofan aircraft engines placed under the wing consoles symmetrically with respect to the plane of symmetry of the aircraft. The technical solution of the aircraft under consideration provides for the use of various engines in aircraft of various groups in terms of passenger capacity and flight range.

The aircraft solution under consideration involves the connection of jet engines with fuel supply lines with a fuel system equipped with two groups of fuel tanks, each of which includes the first and second fuel tanks located in the wing consoles, while the first tanks of the groups are located in the root of the wing consoles, and the second tanks - on their periphery. The fuel tanks of the wing consoles are separated from each other by sealed ribs.

The technical solution of the aircraft under consideration provides, in addition, the possibility of placing an additional supply of fuel in the central fuel tank located in the center section, the side walls of which are formed by the front and rear center section spars and side ribs.

In addition, the aircraft of the family are equipped with a landing gear comprising two main supports and a nose support.

Based on the considered technical solution of the aircraft, it is possible to develop a family of passenger aircraft, the model range of which consists of three groups of aircraft of different passenger capacities (see Belyaev V.V. Passenger airplanes of the world. M: publishing house ASPOL, Argus, 1997, pp. 108-121 ): the first group includes aircraft designed to carry 124 ... 148 passengers (A319-110, A319-130 aircraft), the second group of aircraft includes aircraft designed to carry 164 ... 179 passengers (A320-110, A320-210, A320-230 ), the third group includes aircraft designed for transportation of 200 ... 220 assazhirov (A321-100, A321-200).

The second group contains three aircraft designed for different flight ranges: A320-110 - the shortest - 4900 km, A320-210 - the intermediate - 5185 km and A320-230 - the largest - 5370 km when transporting 150 passengers with baggage. The third group includes three aircraft designed for different flight ranges: A321-100 - the shortest - 4170 km and the intermediate - 4350 km, A321-200 - the longest - 5000 km when carrying 185 passengers with baggage.

The considered technical solution of the family of passenger aircraft is suitable for aircraft designed for passenger capacity from 148 to 220 passengers, however, when switching to aircraft of lower dimensionality, the technical and economic indicators of the family are unsatisfactory. Firstly, the degree of unification of aircraft in the model range is insufficient. Compared to the analogs discussed above (L-1011 and DC-10 family of aircraft), all models of the family of aircraft use a swept wing of the same geometry, however, the family of aircraft of the family includes aircraft with different engines, the structure and layout of the fuel tanks of aircraft of various groups and the maximum estimated flight range.

Secondly, the technical characteristics of the high-speed swept wing and the fuel system do not make it possible to use them as part of the model range of regional passenger airplanes of lower passenger capacity, and consequently, of smaller dimensions. Thus, the wing characteristics are not optimal for use when placing modern turbofan jet engines with significant dimensions near the fuselage under the wing consoles, the mass of the fuel system is large, its stable operation and reliable controllability of the aircraft are not ensured.

The closest analogue to the claimed technical solution of a high-speed swept wing is a wing (see RF patent 2228282, IPC В64С 3/14, publ. 05/10/10/2004), each of whose consoles is composed of six parts, separated from each other by basic sections parallel to the plane of symmetry of the aircraft .

In this technical solution, each of the wing consoles is made up of six parts sequentially placed along each console: the first part from the side section to the GG section (see FIG. 1 of the description to the said patent), the second from the GG section to the FF section, the third from the section FF to section DD, fourth from section DD to section CC, fifth from section CC to BB, sixth from section BB to section AA. Parts of the wing consoles smoothly interfaced with each other, while the wing is formed as a single spatial system based on a non-planar median surface. The profiles of the sections of the wing consoles are made with a change in the position of the midline of the profiles relative to the chords: the position of the middle line of the profiles relative to the chord smoothly changes both in the wing section and in the span of the console. The angle of the geometric twist of the sections and the relative thickness of the profiles of the sections smoothly vary in the scope of the console.

The profiles of the sections of the parts of the wing consoles of this technical solution adjacent to the side section, starting from the first from the side section, are made with sections with an S-shape of the middle lines of the profiles made up of two half-waves. One of the half-waves is made with negative concavity: the midline of the profile is located below the chord. Another half-wave is made with positive concavity: the midline of the profile is located above the chord. The profiles of the sections on the wing periphery adjacent to its tip are made with positive concavity: the middle lines of the profiles along the entire chord are located above the chord.

In the considered technical solution, the half-wave of the S-shaped middle surface with positive concavity is preceded by a half-wave with negative concavity during the transition from the wing tip to its rear part.

In addition, in the transition from the side to the wing end section, the section profiles are made with a monotonic decrease in the angle of the geometric twist of the sections according to the nonlinear law from positive values to negative values and a decrease in the maximum relative thickness of the profiles to 9 ... 10%. In this technical solution, the angle of the geometric twist of the side section is close to 2 degrees, in the end sections, the angle of the geometric twist of the sections is selected from the range from minus 2 to minus 2.5 degrees, and the transition from a positive value of the angle of the geometric twist to negative values is made in the section with a relative coordinate wing span close to 30%. In this technical solution, the maximum relative thickness of the section in the side section slightly exceeds 11%, and in terms of wing span decreases according to a law close to linear.

. В бортовом сечении значение максимальной положительной вогнутости выбрано из диапазона

и размещено по бортовому сечению по координате

. При переходе от бортового сечения к законцовке крыла положение максимального значения положительной вогнутости смещается к хвостику профиля до значения

. Максимальные относительные толщины профилей располагаются при

.In addition, in the considered technical solution of the high-speed swept wing in the side section, the transition from the half-wave of the S-shape of the middle lines with positive concavity to the half-wave with negative concavity was made at a relative coordinate

. In the side section, the value of the maximum positive concavity is selected from the range

and placed along the side section along the coordinate

. When passing from the side section to the wingtip, the position of the maximum value of positive concavity shifts to the tail of the profile to the value

. The maximum relative thicknesses of the profiles are located at

.

, стреловидностью по передней кромке на первых, вторых и третьих частях консолей χ=25…35° и стреловидностью по передней кромке на четвертых, пятых и шестых частях консолей χ=0…30°.In accordance with this technical solution, the considered high-speed swept wing is used for a subsonic aircraft wing with lengthening λ = 9 ... 11, narrowing

, sweep along the leading edge on the first, second and third parts of the consoles χ = 25 ... 35 ° and sweep along the leading edge on the fourth, fifth and sixth parts of the consoles χ = 0 ... 30 °.

The technical solution of the high-speed swept wing does not allow for high aerodynamic characteristics at cruising flight modes at M = 0.75 ... 0.82 using modern engines with a high bypass ratio and in a wide range of lift coefficient С _у , which is typical when using the wing as part of an aircraft with different passenger capacity and with a different flight range. In addition, the technical solution has not resolved the issue related to the placement of large-bypass engines under the wing due to the small distance between the engine nacelle loop and the Earth's surface.

The closest analogue to the fuel system of family aircraft is the technical solution of the fuel system (see IL-86, Vnukovo Production Association, Moscow, 1988, part 1, pages 99-104, part 2, pages 45-62) implemented in the aircraft IL-86.

In this technical solution, the wing contains a transverse power set made of ribs, a longitudinal power set including front and rear spars and stringers, and a casing including its upper and lower parts. The center section, made in the form of a caisson, is bounded by the front and rear spars, side ribs, upper and lower panels and is equipped with longitudinal walls parallel to the side members. Cutouts are made in the walls of the center section.

Each of the wing consoles is made in the form of a caisson, bounded by the front and rear spars, the lower and upper parts of the skin and equipped with ribs, some of which are sealed, and a longitudinal power set. The wing in this technical solution is made with a positive angle of transverse V.

The fuel system comprises a central fuel tank and two groups of fuel tanks. The central fuel tank is located in the center section, it is made in the form of a caisson, bounded by the front and rear spars and side ribs and is divided by the walls of the center section into compartments. Due to cutouts in the walls of the center section, the compartments of the central fuel tank of this technical solution are made communicating with each other.

Two groups of fuel tanks are located in the wing consoles, bounded by the front and rear side members, the lower and upper parts of the wing console skin, and are separated from each other by sealed ribs. Each tank group includes the first fuel tank (in the terminology of the source tanks No. 2 and No. 3) adjacent to the fuselage and separated from the central fuel tank by the side rib, the second fuel tank (in the terminology of the source tanks No. 1A and No. 4A), located on the periphery wing consoles, the third fuel tank (in the terminology of the source tanks No. 1 and 4), located between the first and second fuel tanks. In addition, in the internal volumes of the third fuel tanks of each console there are consumable and main compartments (in the terminology of the source, the main parts of the third tanks). One of the walls of the consumable compartment is aligned with the rib separating the third fuel tank from the first fuel tank.

The considered technical solution assumes the presence in each group of tanks of two consumable compartments connected by fuel supply lines to engines. At the same time, the consumable compartments of the solution in question are located in the internal volumes of the pre-expenditure compartments located in the internal volumes of the first group tanks (in the terminology of the source of tanks No. 2 and No. 3) and in the internal volumes of the third group tanks (in the terminology of the source of tanks No. 1 and No. 4). Consumption compartments in this technical solution are located in the internal volumes of the third fuel tanks near the axis of the wing consoles.

In addition, the fuel system is equipped with fuel transfer lines with transfer pumps, the inputs of which are connected to the internal cavities of the first tank and the main compartment of the third tank of each group, and the internal cavity of the central tank is connected to the inputs of two fuel transfer lines, and the output of one of them is connected to the first , and the other way out - with the second group of tanks.

In the considered technical solution, the exits of the fuel transfer lines from the first tanks are located in the pre-discharge compartments, which are connected to the consumable compartments by additional fuel transfer lines.

In addition, the fuel system of the considered technical solution is made with the possibility of overflow of fuel from the second fuel tanks into the main compartments of the third fuel tanks. This is ensured by the presence of a pipeline between the second tanks and the main compartments of the third tanks, equipped with an electrically controlled valve and non-return valves, which ensure the movement of fuel by gravity from the second tanks.

The considered technical solution is used in the wing with a positive angle of the transverse V.

In the process of fuel production, the central fuel tank is first emptied, then the first fuel system tanks and the third tanks begin to empty simultaneously. By emptying the third tanks, continuing to generate fuel from the first tanks of the fuel system, the emptying of the second tanks of the system begins. After emptying the first and second tanks, fuel is last generated from the supply compartments.

The technical solution of the fuel system under consideration can be mainly used for wide-body aircraft of large passenger capacity and flight range. The use of this solution for aircraft of lower passenger capacity and flight range, and therefore designed for smaller volumes of refueling fuel, has several disadvantages.

The relative mass of the fuel system in accordance with this technical solution is large due to the complex structure of the fuel system, including pre-use compartments and additional fuel transfer lines. Large and undeveloped fuel residues in the fuel system. The disadvantage of this system is the complexity of controlling the production of fuel from the tanks of the system, which requires the presence of special systems in the fuel system to control fuel production.

In addition, this fuel system does not provide the possibility of its use for a family of passenger aircraft, consisting of groups of aircraft of different passenger capacities, each of which includes aircraft with different flight ranges. Aircraft of different passenger capacity and flight range differ in the longitudinal dimensions of the fuselage and different masses of refueling fuel for aircraft of different flight ranges. When using the specified technical solution of the fuel system for a family of aircraft, the longitudinal alignment of the aircraft changes, which greatly complicates the process of controlling the aircraft.

The technical problem solved by the claimed aircraft solution is the development of an aircraft with high technical, economic and technical and operational indicators, including a high degree of unification, mass excellence, reliable controllability, high safety, combined with the possibility of creating on its basis a model range of aircraft with different passenger capacities and flight range.

The technical problem solved by the claimed solution of the high-speed swept wing is the development of a wing with high aerodynamic characteristics, providing the possibility of its use without changing the geometric parameters in the aircraft, intended for use in the family of passenger aircraft of different passenger capacity and flight range in combination with the possibility of placing two engines under the wing with large dimensions for use in the range of cruising flight modes, I characterize REGARD number M = 0.75 ... 0.82.

The technical problem solved by the claimed aircraft fuel system is the development of a fuel system device that provides a reduction in the mass of the fuel system and a reduction in unreleased fuel residues, improved controllability of the aircraft and increased uptime of its operation in emergency situations, combined with the possibility of its use in a family of aircraft, including groups of aircraft with different passenger capacity and different maximum estimated range.

The claimed technical solution of the aircraft, the technical task is solved as follows.

The technical solution of the aircraft, made with the possibility of changing the maximum estimated range, is known. The fuselage of the aircraft consists of sequentially placed nose, front, center, rear and tail compartments. In a known solution, the front and rear compartments are made with a length that enables the placement of a different number of passengers on board the aircraft. The chassis includes two main and bow supports.

The high-speed swept wing of the known aircraft solution consists of a center section and wing consoles, includes longitudinal and transverse power set and casing with its lower and upper parts.

The wing is made with a positive angle to the transverse V.

The front and rear spars are included in the longitudinal power set of the known wing, and ribs, some of which are sealed, are included in the transverse power set. Each wing console is equipped with a “dry compartment”, one of the walls of which is combined with the front side member.

Consoles of the known wing are made with a decrease in the maximum relative thickness of their profiles during the transition from the side section of the console to the tip, and sections with a local minimum of the maximum relative thickness of the profiles are placed on the half of the consoles adjacent to the side sections.

The aircraft of the known solution is equipped with a power plant containing two turbofan jet engines mounted under the wing consoles symmetrically relative to the plane of symmetry of the aircraft. Engines are connected by fuel supply lines to the fuel system.

The fuel system of the known solution is equipped with two groups of fuel tanks. Fuel tanks are separated from each other by sealed ribs. Each of the groups of tanks includes the first and second fuel tanks located in the wing consoles, with the first tanks of the groups placed in the root of the wing consoles, and the second tanks on their periphery. In addition, in the known aircraft solution, its fuel system is equipped with a central fuel tank located in the center section, the side walls of which are formed by the front and rear wing spars and side ribs.

In the claimed technical solution of the aircraft new is that the installation plane of turbofan aircraft engines are located near one third of the wing span. During the transition along the wing console from the side section to one quarter of the wing half-span, the swept wing of the aircraft is made with an increase in the angle of the transverse V wing, and with a further transition to the wing tip, with its decrease. In the claimed solution, profiles from the side section to at least half of the wing span are made with positive angles of geometric twist of their sections.

The local minimum of the maximum relative thickness of the profiles of the wing consoles of the proposed solution is shifted compared to the analog to one third of the wing span.

In each of the groups of tanks of the fuel system of the claimed aircraft solution, a third fuel tank is additionally included, located in the wing consoles between the first and second tanks of the groups. The third tanks are divided into main and consumable compartments. Consumption compartments are connected to the fuel supply lines. One of the walls of the consumable compartments is combined with a sealed rib separating the third fuel tanks from the first, and the other with the rear spar.

The fuel system of the claimed aircraft is made possible to produce fuel first from the central fuel tank, then from the first tanks, then from the second tanks, and last of all from the third tanks of the groups. The total volume of two groups of tanks and the volume of the entire central fuel tank are selected based on the assurance of fueling the aircraft with the highest value of the maximum estimated flight range.

Dry compartments of the fuel system of the claimed aircraft solution are located on both sides of the sealed ribs that separate the first fuel tanks from the third. In addition, a closed U-shaped stringer tightly connected to the upper part of the wing skin is included in the said longitudinal power set of each console. The internal cavity of the U-shaped profile is made communicating with the first and second tanks of the fuel system. U-shaped stringers for most of their length are located near areas with the maximum relative thickness of the profiles of the airplane wing consoles.

In the claimed decision of the aircraft, in addition, it is new that the nose and rear compartments of the fuselage can be made with a length that provides the ability to transport aircraft from 58 to 68, from 66 to 83 and from 83 to 108 passengers.

In addition, in the claimed decision of the aircraft, the equivalent diameter of the fuselage can be selected from the range of 3.3 ... 3.55 m, while the elongation of the fuselage of the aircraft, designed to carry from 58 to 68 passengers, can be selected from the range of 6.5 ... 6, 7, designed to carry from 66 to 83 passengers - from the range of 7.3 ... 7.6, and the elongation of the fuselage of the aircraft, designed to carry from 83 to 108 passengers, can be selected from the range of 8.2 ... 8.6.

The fuselage and high-speed swept wing of the claimed aircraft can be made to provide the possibility of flying at a cruising speed with a Mach number from 0.75 to 0.82.

In addition, in the claimed decision of the aircraft in the center section can be installed two cut-out walls parallel to the side members, while the walls of the central fuel tank can be divided into three compartments, the first of which is located at the rear side member, the third - at the front side member, and the second - between the walls. In this case, cutouts in the walls of the center section can be made with the possibility of hermetic overlapping of their covers. The total volume of two groups of tanks and the volume of the first compartment of the central fuel tank can be selected from the calculation of providing fueling aircraft with the lowest value of the maximum estimated range. The total volume of two groups of tanks and the volume of the first and second compartments of the central fuel tank can be selected from the calculation of providing fuel for the aircraft with an intermediate value of the maximum estimated flight range.

In addition, in the claimed decision of the aircraft, a sealed rib separating the first and third tanks of the wing consoles, at least in its tail section, can be placed within the first quarter of the wing half-span.

High technical and economic performance of the aircraft is achieved, first of all, due to the development of a high-speed swept wing of a uniform geometry for the entire model range of aircraft with different fuselage lengths and ensuring the placement of large-size jet engines under the wing consoles.

The proposed aircraft solution, which provides for the installation of engines and the placement of a local minimum of the maximum thickness of the profiles of the wing consoles near one third of the wing half-span, combined with an increase in the angle of the transverse V wing to one quarter of the half-span of the wing, followed by its reduction and execution of profiles with positive angles of geometric twist to at least half wing span provides high aerodynamic performance of the aircraft when placing engines significant dimensions under wing consoles and changing overall dimensions of the fuselage.

The inclusion in the fuel system of the aircraft of the third tanks located in the wing consoles between the first and second tanks, and the inclusion in the longitudinal power set of each console of the stringer of a closed U-shaped profile connecting the cavities of the first and second tanks of the fuel system, makes it possible to generate fuel first from the central fuel aircraft tank, then from the first tanks, then from the second tanks, and last of all from the third tanks of the groups, which not only positively affects the stability and controllability of the aircraft, but It makes it possible to vary the maximum refueling aircraft in which the total volume of the two groups of cans and the total volume of the central fuel tank can provide a refueling aircraft with the highest maximum design range.

Combining one of the walls of the consumable compartments with a sealed rib separating the third fuel tanks from the first and the other with the rear spar, placing dry compartments in each wing console at the front spar on both sides of the sealed ribs separating the first fuel tanks from the third, and turning on in the longitudinal power set of the stringer of a closed U-shaped profile, the mass of the wing consoles is reduced. Moreover, the placement of stringers of a closed U-shaped profile near areas with a maximum relative thickness of profiles additionally ensures reliable operation of the fuel system. Design estimates show that the mass of the aircraft structure in accordance with the claimed solution is reduced by 5 ... 7% compared with analogues.

In addition, the location of the consumable compartment, the fuel from which is selected last, relative to the “dry” compartment of the console increases flight safety, since the consumable compartment is maximally distant from the zone of non-localized expansion of the rotors of the main propulsion system.

Features of changing the angle of the transverse V wing and the angle of the geometric twist of the sections along the wing console, in combination with the possibility of placing a sealed rib separating the first and third tanks of the wing consoles, helps to increase the safety and reliability of the fuel system in emergency situations.

As in the closest analogue, the transition from one group of aircraft in terms of passenger capacity to another is provided by changing the length of the fuselage, for which the fuselage of each of the aircraft of the family is divided by technological joints into sequentially placed bow, front, center, rear and tail compartments, with the bow, center and tail the compartments of the aircraft of the family are the same for all aircraft of the family, and the front and rear compartments are made with a length that provides the possibility of placing on board the aircraft number passengers corresponding to the group of aircraft in terms of passenger capacity.

A change in the maximum flight range of aircraft models within a family group can be achieved in the claimed solution by changing the volume of the central fuel tank available for refueling. In this case, cutouts in the walls of the center section can be made with the possibility of hermetic overlapping of their covers. In this case, the total volume of two groups of tanks of the fuel system and the volume of the first compartment of the central fuel tank can be selected from the calculation of fueling an aircraft of any group with the lowest flight range in its group, the total volume of two groups of tanks and the volume of the first and second compartments of the central fuel tank can be selected on the basis of providing fueling for an aircraft of any group with an intermediate value of the flight range in its group, and the total volume of two groups of tanks and the volume of all prices eral fuel tank may be selected from the software calculating refueling aircraft of any group with the highest value of the flight distance in the group.

The possibility of overlapping cutouts of the walls of the center section with covers helps to shift the center of gravity of the aircraft to the rear centering area, reduce undeveloped fuel residues, and reduce the dynamic effect of fuel on the aircraft structure. These factors are especially relevant for shortest and intermediate range aircraft.

Other technical solutions are possible, providing the ability to change the mass of refueling fuel for the maximum estimated range of the aircraft.

The inventive device of family aircraft together with the indicated levels of the maximum estimated flight range can provide another, fourth, level of flight range, which is provided by the fuel supply, placed only in the tanks of the wing consoles. In this case, the central fuel tank located in the center section does not refuel, and structurally it is insulated by hydraulic connections with the fuel tanks of the wing consoles.

Design studies show that, based on the proposed decision of the aircraft, a model range of passenger aircraft can be developed, including three groups of aircraft designed to carry from 58 to 68, from 66 to 83 and from 83 to 108 passengers. With a round cross-sectional shape of the fuselage with an equivalent diameter of 3.3 ... 3.55 m, the elongation of the fuselage of the aircraft can vary from 7.3 to 8.6. With a wing span of 25 ... 30 m, in each group of aircraft in terms of passenger capacity, three aircraft models with the smallest, intermediate and largest maximum estimated range of flight in their group can be represented. In this case, each group of aircraft in terms of passenger capacity can include planes with the smallest (3100 ... 3300 km), intermediate (3900 ... 4100 km) and the largest (4600 ... 4900) maximum estimated flight ranges.

Thus, the above characteristics of the proposed technical solution of the aircraft allow you to choose its device, which is characterized by high technical and operational performance, ensuring the development of a model range of passenger aircraft with different passenger capacity and flight range, designed primarily to serve regional lines. At the same time, the proposed aircraft is characterized by a high degree of unification, a decrease in comparison with the analogues of the mass of the aircraft, reliable controllability and safety.

The technical task of the claimed solution to the high-speed swept wing is solved as follows.

A technical solution is known for a high-speed swept wing, consisting of a center section and consoles. Each of the wing consoles is composed of six parts delimited from each other by sections parallel to the longitudinal plane of symmetry of the aircraft, placed sequentially along the span of the wing consoles and smoothly mated with each other. Consoles are formed as a single spatial system based on a non-planar median surface.

In the known solution, the profiles of the sections of the consoles adjacent to the fuselage are made with an S-shaped center line composed of two half-waves, one of which is made with a negative concavity, and the other with a positive one. Section profiles on the periphery of the wing consoles are made with positive concavity of the midlines.

In the known solution of the wing, in addition, when switching from the side to the wing end section, the profiles are made with a monotonic decrease in the angle of geometric twist from positive to negative values and a decrease in the maximum relative thickness of the profiles to a value of 9 ... 11%.

In the claimed decision of the high-speed swept wing, it is new that profiles with an S-shape of the midline are placed on the first five, counting from the fuselage, parts of each console. In addition, during the transition along the chord of the profile from its toe to its tip, a half-wave with negative concavity of the midline of the profiles is preceded by a half-wave with positive concavity. Profiles with positive concavity of the midline are placed on the sixth of each of the consoles. In the claimed solution, a new one is that on the first, fourth and fifth parts of each of the wing consoles, two S-shaped half-waves of the midline of the profiles are preceded by an additional half-wave with a positive concavity of the midline of the profiles.

In the claimed decision of the wing, in addition, near the side section, the value of the angle of the transverse V wing along the leading edge is selected from the range of 6 ... 8 degrees, and when switching from the side section to the border between the second and third parts of the consoles, the wing is made with increasing angle of the transverse V wing along leading edge to values from 9 to 12 °, and upon further transition to the wingtip, the wing is made with decreasing the angle of the transverse V wing to a value from the range 5 ... 6 °.

In addition, near the side section, the value of the angle of the geometric twist is selected from the range of plus 3 ... plus 3.8 degrees, and the value of the maximum relative thickness of the profiles is selected from the range of 14 ... 16%. In the claimed solution, at the border between the third and fourth parts of the wing consoles, a local minimum of the maximum relative thickness of the wing profiles is placed, and when moving along the fourth part from the border with the third part to the border with the fifth part, the profiles are made with an increase in the maximum relative thickness of the profiles.

In the claimed solution of the high-speed swept wing, the transition from sections with positive values of the angle of geometric twist of the profiles to sections with negative values of the geometric twist of the profiles can be performed on the border between the fifth and sixth parts of the wing consoles.

, где

- относительная координата по хорде профиля.In addition, in the claimed decision of the wing, the maximum relative thicknesses of the profiles should be located at

where

- relative coordinate along the profile chord.

и при переходе от бортового сечения к границе между пятой и шестой частями консолей крыла значение максимальной относительной вогнутости профилей целесообразно уменьшать от значений из диапазона

до 0, где

- относительная координата по хорде профиля.The area with the maximum negative concavity of the profiles in the inventive high-speed swept wing can be located at

and when passing from the side section to the boundary between the fifth and sixth parts of the wing consoles, it is advisable to reduce the value of the maximum relative concavity of the profiles from values from the range

to 0, where

- relative coordinate along the profile chord.

в диапазон

, где

- относительная координата по хорде профиля.The relative coordinate of the transition from a half-wave with negative concavity of the S-shaped middle line of the profiles to a half-wave with a positive concavity of the middle line of the profiles when moving from the side section to the boundary between the fifth and sixth parts of the wing consoles in the claimed solution is advantageously shifted from the range

in range

where

- relative coordinate along the profile chord.

, при этом при переходе от бортового сечения к законцовке крыла значение максимальной относительной вогнутости профилей целесообразно увеличивать от

до

, где

- относительная координата по хорде профиля.The section of maximum positive concavity of the profiles should be located at

, in this case, when passing from the side section to the wingtip, it is advisable to increase the value of the maximum relative concavity of the profiles from

before

where

- relative coordinate along the profile chord.

Most preferably, the inventive high-speed swept wing is performed with a sweep angle along the leading edge on the first, second and third parts of the wing consoles selected from the range 31 ... 33 °, and on the fourth, fifth and sixth parts of the wing consoles - from the range 26 ... 28 °. Moreover, the trailing edge of the wing on the first, second, third and fourth parts of the wing consoles can be made without sweep, and the angle of sweep of the trailing edge on the fifth and sixth parts of the wing consoles can be selected from the range of 20 ... 23 °.

, координаты границы между первой и второй частями консолей крыла - из диапазона

, второй и третьей частями -

, третьей и четвертой частями -

, четвертой и пятой частями -

, пятой и шестой частями

от полуразмаха крыла.When using the claimed solution of a high-speed swept wing in a regional passenger plane, it is most preferable to choose the coordinates of the wing cross sections

, the coordinates of the boundary between the first and second parts of the wing consoles are from the range

, second and third parts -

, the third and fourth parts -

, fourth and fifth parts -

fifth and sixth parts

from the wing span.

The set of features of the inventive high-speed swept wing allows you to solve the technical problem.

Calculations show that the use of a high-speed swept wing with the indicated character of the median surface provides a close to optimal flow around the wing with engines located under the wing in the engine nacelles and close to optimal pressure distribution on the aerodynamic bearing surface. Moreover, the nature of the median surface, characterized by the presence of profiles with a negative concavity of the midline on a large part of the wing console adjacent to the fuselage, significantly reduces harmful interference between the wing and the fuselage.

The presence of additional half-waves with positive curvature on the nose of the profile in the areas bordering the zone with the placed engine, in combination with the selected laws of changing the angle of the transverse V wing and changing the maximum relative thickness of the profiles along the half-span of the wing, form not only favorable conditions for the flow of the nacelle with the engine under the wing consoles, but also provide optimal conditions for placement under the wing consoles of a significant-sized engine. This increases the distance of the engine from the runway.

In addition, the values of the angles of the transverse V wing along the leading edge provide the necessary lateral stability characteristics.

In addition, the adopted nature of the change in the angle of the geometric twist of the profile, increases the aerodynamic quality of the wing. Placing the maximum relative profile thickness in the range from 35 to 42% of the chord length provides an additional increase of 2 ... 4% in the value of aerodynamic quality.

до нуля дополнительно снижает значение аэродинамического сопротивления на 3…5%.Placing sites with maximum values of negative relative concavity of the profiles on the parts of the wing consoles from first to fifth in the range of relative coordinates along the chord length from 35 to 42% and their smooth change from the value

to zero additionally reduces the value of aerodynamic drag by 3 ... 5%.

Calculations show that the aerodynamic drag of the wing decreases by about the same amount due to the shift of the transition from the negative half-wave of the S-shape of the midline of the profile to the positive half-wave from the range along the relative coordinate along the length of the chord 75 ... 80% to the range 45 ... 50% to the boundary between the fifth and sixth parts of the aerodynamic bearing surface.

в бортовом сечении до значения

в концевых профилях консоли повышает аэродинамическое качество на 5…8%.In addition, the placement of sections with the maximum relative positive concavity of the S-shaped middle profile lines along the wing console in the range from 80 to 90% of the profile chord length and its increase from the value

in the side section to the value

in the end profiles of the console increases the aerodynamic quality by 5 ... 8%.

The presence of sagging on the first and fourth parts of the wing consoles along the front and rear edges optimizes the flow regime in the area adjacent to the fuselage, and allows you to increase the mass of fuel in the fuel tanks that can be placed in the wing consoles. The presence of a direct influx along the trailing edge additionally makes the operation of wing mechanization more efficient in takeoff and landing modes.

The proposed boundaries between the parts of the bearing aerodynamic surface optimize the flow regimes of the wing and make it possible to place a significantly larger engine under the wing consoles along the boundary between the third and fourth parts of the aerodynamic bearing surface.

, в сочетании с мотогондолами двигателей, размещенными под консолями крыла, обеспечивает высокие значения аэродинамического качества на уровне К_макс=15…17 в широком диапазоне чисел Маха, при этом максимальное значение показателя К_макс*М на уровне 12…13 достигается при М=0,76…0,82, что говорит о высоких аэродинамических характеристиках скоростного стреловидного крыла.Calculations and tests show that the claimed solution, for example, when used in a wing with a span of 26 to 30 m with an elongation of λ = 9 ... 11, a narrowing not less than

, in combination with engine nacelles located under the wing consoles, provides high values of aerodynamic quality at the level of K _max = 15 ... 17 in a wide range of Mach numbers, while the maximum value of K _max * M at 12 ... 13 is achieved at M = 0 , 76 ... 0.82, which indicates the high aerodynamic characteristics of the high-speed swept wing.

In addition, the aerodynamic characteristics of the wing with the engine nacelles under the wing remain stable when the mass and length of the fuselage of the aircraft change: for example, the aerodynamic quality of the wing varies within no more than 1 ... 4% when switching from an airplane designed to carry 60 passengers to an airplane designed to transportation of 95 passengers. This makes it possible to use the inventive high-speed swept wing without changing geometric parameters for a family of passenger aircraft of various passenger capacities and flight ranges in combination with the placement of two engines with a large bypass ratio in the range of cruising flight modes, characterized by the number M = 0.75 ... 0.82.

The inventive fuel system of the technical problem is solved as follows.

In the well-known technical solution of the fuel system, fuel tanks are located in the center section and wing consoles. The wing of the known technical solution is made with a positive angle of the transverse V and includes the front and rear spars, the casing with its upper and lower parts, ribs, some of which are sealed, a longitudinal power set of wing consoles, including stringers. In addition, in a known technical solution, the center section includes cut-out walls parallel to the side members.

In a known technical solution, the fuel system is equipped with a central fuel tank located in the center section, made in the form of a caisson, bounded by front and rear spars and side ribs and divided by the walls of the center section into compartments. In addition, in the known technical solution, the fuel system is equipped with two groups of fuel tanks located in the wing consoles, bounded by the front and rear spars, the lower and upper parts of the casing and separated from each other by sealed ribs. In this case, each group of tanks includes a first fuel tank adjacent to the fuselage and separated from the central fuel tank by an onboard rib, a second fuel tank located on the periphery of the wing consoles, and a third fuel tank located between the first and second fuel tanks. At the same time, the expendable and main compartments are allocated in the third fuel tank. The consumable compartment is located in the internal volume of the third fuel tank, while one of its walls is aligned with the rib separating the third fuel tank from the first fuel tank. In addition, each group of tanks in the known solution is equipped with a fuel supply line to the engine, the input of which is connected to the consumable compartment. In the known technical solution, in addition, the fuel system is equipped with fuel transfer lines with transfer pumps, the inputs of which are connected to the internal cavities of the first tanks and the main compartments of the groups of tanks, and the internal cavity of the central tank is connected to the inputs of two transfer lines, the output of one of them being connected with the first, and the other way out - with the second group of tanks. In addition, the ribs separating the second tanks from the main compartments of the third tanks are equipped with check valves located near the lower wing casing and made with the possibility of moving fuel from the second tanks to the main compartments of the third tanks,

In the claimed technical solution, it is new that the second wall of the consumable compartment is combined with the wall of the rear spar. In addition, in the claimed technical solution, the new is that in the upper parts of the walls separating the consumable compartments from the main compartments of the third fuel tanks, and in the upper parts of the ribs separating the third fuel tanks from the second, slots are made, and in the longitudinal power set of each console a closed U-shaped stringer is included, hermetically connected to the upper part of the wing skin. The internal cavity of this stringer is connected to the internal cavities of the first and second fuel tanks.

In addition, in the claimed solution, each group of fuel tanks is equipped with overflow channels providing the possibility of fuel overflow from the first fuel tanks to the central fuel tank, the inputs of which are located in the upper parts of the first fuel tanks, and the outputs are made with the possibility of fuel overflow in one of the compartments of the central a fuel tank, wherein the exits of the said fuel transfer lines from the first tanks and the central fuel tank are connected to the third tanks of the groups, and the exits of the transfer lines and from the main compartments of the third tanks are connected to the consumable compartments of the third tanks.

In the claimed solution, each group of fuel tanks can be equipped with a fuel transfer line from the second tank to the third.

In addition, in the claimed solution of the fuel system, the exits of the said fuel transfer lines from the first and second group tanks and the central fuel tank can be connected to the main compartments of the third group tanks.

In addition, in the claimed solution of the fuel system, the exits of the said fuel transfer lines from the first and second group tanks and the central fuel tank can be connected to the consumable compartments of the third group tanks.

In addition, the new solution in the claimed solution is that in the lower parts of the walls separating the consumable compartments from the main compartments of the third fuel tanks, and ribs separating the first fuel tanks from the main compartments of the third tanks, check valves can be placed, providing fuel overflow from the first fuel tanks to the main compartments of the third tanks and from the main compartments of the third fuel tanks to consumable compartments and the exclusion of fuel overflow in the opposite direction.

In addition, in the claimed solution, it is new that the said overflow channels can be equipped with check valves, providing the possibility of overflow of fuel from the first tanks into the central compartment and elimination of overflow of fuel in the opposite direction.

In addition, in the claimed solution, cutouts of the walls of the center section can be made with the possibility of hermetic overlapping of their covers.

The claimed features of the fuel system make it possible to solve the technical problem formulated above.

The implementation of the fuel system in accordance with the claimed solution allows to reduce the mass of the fuel system by 10-15%, and the mass of undeveloped fuel residues from 2 ... 3% to 0.5 ... 1%.

The presence of slots in the upper parts of the walls separating the consumable compartments from the main compartments of the third fuel tanks, the ribs separating the third fuel tanks from the second tanks, overflow channels connecting the first fuel tanks and the central fuel tank, closed U-shaped stringers, the internal cavities of which connect the second and first fuel tanks, and check valves in the ribs separating the second tanks from the third fuel tanks, provides a given sequence of fuel generation from the tanks of the fuel system Without the use of complex control systems for the generation of fuel from the tanks of the fuel system, which reduces the weight of the system.

In addition, the placement of the consumable compartments in the tail of the wing by combining the walls of the consumable compartments with the rear side member not only reduces the mass of undeveloped fuel residues, but also improves the alignment of the aircraft.

The presence of check valves in the lower parts of the walls separating the supply tanks and the third tanks of the fuel system, ribs separating the third fuel tanks from the first and second tanks of the fuel system, as well as in overflow channels connecting the first tanks and the central fuel tank, preventing unplanned fuel overflows between tanks of the system, additionally contribute to increasing the reliability of the fuel system in case of accidental unplanned evolution of the aircraft. The introduction into the fuel system of the lines for pumping fuel from the second fuel tanks to the third fuel tanks does not change the specified sequence of fuel generation from the fuel system tanks, however, delaying the time of fuel production from the main compartment of the third fuel tank expands the ability to control the change in the position of the center of gravity of the aircraft in flight.

The connection of the fuel transfer lines from the second and first fuel tanks and the central tank of the fuel system with the main compartments of the third tanks in combination with the presence of the lines for pumping fuel from the main compartments to the consumable compartments of the fuel tanks increases the reliability of the fuel system: with this pumping scheme, the fuel settles in the main compartments while gas inclusions in the fuel are separated from the liquid phase of the fuel.

The connection of the fuel transfer lines from the second and first fuel tanks, the main compartments of the third tanks and the central tank of the fuel system directly to the consumable compartments of the third tanks makes it possible to simplify the fuel system circuit and slightly reduce its weight.

The presence of check valves in the overflow channels connecting the first tanks of the fuel system to the central tank prevents overflows of fuel from the central tank to the first tanks of the fuel system during abnormal operating conditions.

In accordance with the claimed decision of the fuel system, cutouts in the walls of the center section can be covered with covers. This makes it possible to manufacture aircraft with a fuel system designed for various levels of maximum fueling. In the manufacture of an aircraft with a short maximum range, one of the walls of the center section can be provided with covers and part of the central fuel tank can be allocated for fueling. However, the other part of the central fuel tank remains “dry”. On the contrary, in the manufacture of an aircraft with the longest flight range, without providing cutouts to the walls of the center section of the lids, it is possible to use the entire central fuel tank for refueling. This improves the alignment of the shortest and intermediate range aircraft. In addition, when using the shortest and intermediate flight ranges for airplanes, disturbances caused by fuel fluctuations in the central fuel tank are reduced by reducing the area of the fuel mirror in the tank.

The claimed group of inventions is illustrated by the following materials:

Figure 1 - General view of the aircraft;

Figure 2 is a technological diagram of the division of the aircraft;

figure 3 - layout diagrams of aircraft of three groups in terms of passenger capacity (side view);

figure 4 - layout diagrams of aircraft of three groups in terms of passenger capacity (top view);

5 is an external view of the aircraft (side view and front view);

6 is a structural power diagram of the wing;

Fig.7 is a typical leaky wing rib (view BB from Fig.6);

Fig.8 is a cross section of a stringer U-shaped profile (view I from Fig.7);

Fig.9 is a cross section of the ribs along LL with Fig.7;

figure 10 is a cross section of the center section of the wing (section II from figure 6);

11 - the wall of the center section with cut-outs (view G-G from Fig.6);

12 is a view of the center section wall with cutouts overlapped by covers;

Fig. 13 is a sectional view of the lid (section along DD in Fig. 12);

Fig - graph of the change in the angle of the transverse V on the wing console;

Fig - graph of changes in the maximum relative thickness of the profile along the wing console;

Fig is a graph of changes in the angle of the geometric twist of the sections along the wing console;

Fig is a General schematic diagram of a fuel system;

Fig. 18 is a diagram of a third tank of a fuel system;

Fig.19-20 is a sequence of fuel generation from the tanks of the fuel system;

Fig - an approximate layout of passenger aircraft, designed to carry 60, 75 and 95 passengers with a step between rows of seats 812.8 mm;

Fig.22 is a cross section of the passenger cabin of the aircraft (section ZZ with Fig.21);

23, 24, 25 — dependency graphs; estimated range — commercial load of the aircraft for 3 groups of aircraft different in passenger capacity;

Fig is a diagram of the wing console;

Fig.27 is a graph of the midline of the profile in the side section (section H-N of Fig.26),

Fig.28 is a graph of the midline of the profile in section G-G of Fig.26,

Fig.29 is a graph of the midline of the profile in cross section F-F with Fig.26,

Fig.30 is a graph of the midline of the profile in cross section D-D of Fig.26,

Fig.31 is a graph of the midline of the profile in section CC with Fig.26,

Fig. 32 is a graph of the midline of the profile in section BB in Fig. 26,

Fig.33 is a graph of the midline of the profile in section AA from Fig.26;

Fig.34 - calculated lines flowing around the upper surface of the wing console;

Fig - estimated distribution of isobars on the upper surface of the wing console;

Fig is a graph of changes in maximum aerodynamic quality (K _max ) and indicator K _max * M for an aircraft, designed to carry 83 ... 108 passengers;

Fig. 37 is a graph of the aerodynamic quality of a wing with engine nacelles for aircraft of various passenger capacities;

Fig. 38 is a comparison of the aerodynamic quality and the indicator K _max * M of the wing of the claimed wing with one of the analogues;

Fig. 39 is a schematic diagram of a fuel system with fuel transfer from the first tanks and the central fuel tank to the main compartments of the third tanks;

40 is a schematic diagram of a fuel system with pumping fuel from the first and second tanks and the central fuel tank to the consumable compartments of the third tanks;

Fig is a schematic diagram of a fuel system with pumping fuel from the first and second tanks and the Central fuel tank to the main compartments of the third tanks;

Fig. 42 is a half-circuit diagram of a fuel system (in accordance with the circuit diagram shown in Fig. 40);

Fig. 43 is a longitudinal sectional view of a wing console along a U-shaped stringer;

Fig.44 is a view of a sealed rib separating the tanks from each other (view MF from Fig.6);

Fig. 45 is a diagram of a central fuel tank with placement of overflow channel outlets in a first compartment of a central fuel tank;

Fig. 46 is a diagram of a central fuel tank with placement of overflow channel outlets in a second compartment of a central fuel tank;

Figures 47-52 are a sequence diagram of the main steps for generating fuel from a fuel system;

Fig.53-54 is a sequence diagram of the main stages of fuel generation from a fuel system equipped with highways for pumping fuel from second tanks to third tanks.

The claimed technical solution of the aircraft is arranged as follows. The aircraft is equipped (see Fig. 1) with a fuselage 1, a high-speed swept wing 2, a power plant with turbofan air-jet engines 3, a fuel system, a landing gear with two main supports 4 and a nose support 5, vertical 6 and horizontal 7 plumage.

The fuselage 1 is divided by technological joints (see figure 2) into sequentially placed nose 8, front 9, central 10, rear 11, tail 12 compartments and the auxiliary power unit 13. In the direction of flight, nose cone 14 is adjacent to the nose compartment 8. Compartments the fuselage is divided by floor 15 into the passenger compartment 16 and the luggage and technical compartments 17.

The bow 8, central 10 and tail 12 of the aircraft fuselage compartments are identical for all aircraft of the model range. The front 9 and rear 11 compartments are made with a length that provides the possibility of placing on board the number of passengers of the corresponding group of aircraft according to passenger capacity. As an example, figure 3-4 presents three aircraft with different passenger capacity (58 ... 68, 66 ... 83 and 83 ... 108 passengers), representing three groups of aircraft of the family. At the same time, when switching from an aircraft designed to carry from 58 to 68 passengers to an aircraft designed to carry from 66 to 83 passengers, the length of the front compartment increases by 1, the rear by 1.5 m, and when switching from an aircraft designed to carry from 58 to 68 passengers, for an aircraft designed to carry from 83 to 108 passengers, the lengths of these compartments increase by 3 m.

Aircraft of the family are equipped with a landing gear that includes two main bearings and a nose support, while it is most advisable to choose a landing gear that is the same for the entire model range of aircraft in the family. The vertical and horizontal plumage is also advisable to choose the same for all aircraft of the family.

The proposed technical solution of the aircraft involves the use (see figure 4) identical in geometry to the speed swept wing 2 for aircraft with different passenger capacities. The wing consists of a center section 18 and wing consoles 19. The wing of the aircraft is made with a positive angle of transverse V (see figure 5, 14).

The wing of the aircraft of the proposed solution (see Fig.6) is equipped with a longitudinal and transverse power set. The front 20 and rear 21 spars and stringers 29 are included in the longitudinal power set, and the ribs 22 are included in the transverse power set, some of which 24, 25, 26, 27 are airtight.

In addition, each wing console is equipped with a “dry” compartment 23, which is not fueled. The “dry” compartment adjoins the front side member: one of its walls is aligned with the front side member 20, while the “dry” compartment is located on both sides of the rib 24. The rib 24 is sealed. The wing ribs are perpendicular to the axis of the wing consoles, with the exception of the side ribs 25 and one rib 28 in each wing console following the side rib 25.

Stringers 29 are included in the longitudinal power set of the wing consoles (see Figs. 7, 8). In addition to these stringers 29, a closed U-shaped profile stringer 30 is included in the longitudinal power set of the wing consoles. This stringer is hermetically connected to the upper part of the wing skin 31.

The wing center section (see FIGS. 6, 10) is formed by the front 20 and rear 21 side members, side ribs 25, upper 32 and lower 33 panels. In addition, in the center section there are two longitudinal walls 34 arranged parallel to the side members. The walls of the center section 34 are provided with cutouts 35. By the walls 34, the internal volume of the center section is divided into three compartments: the first 36 of which are located at the rear side member, the third 38 - at the front side member, and the second 37 - between the walls of the center section 34. The cutouts 35 in the walls 34 of the center section can be made so that they can be sealed by their covers 39 (see Fig. 12, 13).

The high-speed swept wing of the aircraft is made of the same geometry with the top 31 and bottom 40 skins for the entire model range.

The wing of the claimed aircraft is made with a change in the angle of the transverse V along the console (see Fig. 14). In this case, the angle of the transverse V both along the leading edge (curve 41 in Fig. 14) and along the trailing edge (curve 42 in Fig. 14) increases when moving along the wing console from the side section to one quarter of the wing half-span (25%), with further transition to the wingtip - decreases.

In addition, the wing consoles are made with a decrease in the maximum relative thickness of the profiles during the transition from the root section of the console to the tip (see Fig. 15), and a portion with a local minimum of 43 of the maximum relative thickness of the surface profiles is placed on its half adjacent to the side section. The wing of the aircraft in the claimed solution is made to achieve a local minimum of the maximum relative thickness of the surface profiles near one third of the half-span of the wing.

In addition, the profiles of the wing consoles of the claimed aircraft are made with positive angles of geometric twist from the side section to at least half of the wing span (see Fig. 16).

, где

- относительная координата по хорде профиля.It is advisable to omit said closed U-shaped stringer 30 along the profiles of the wing consoles near the location of their maximum relative thickness. For a family of regional aircraft designed to carry from 53 to 108 passengers, it is advisable to place this profile when

where

- relative coordinate along the profile chord.

The most appropriate high-speed swept wing of the proposed aircraft to perform in accordance with the claimed decision of the high-speed swept wing, the description of which is given below.

The power plant contains two turbofan jet engines 3 located in the engine nacelles under the wing consoles symmetrically relative to the plane of symmetry of the aircraft 67. In the claimed solution, it is advisable to equip all aircraft of the family with the same turbofan air-jet engines. So, when using the proposed solution to create a family of regional passenger airplanes, the indicated wing features allow placing under its consoles jet engines of a large degree of double-circuit and large dimensions, which makes it possible to use them without changes in all aircraft of the family of the family. The jet engine installation planes should be located near one third of the wing half-span, which, in combination with the noted features of the high-speed swept wing of the claimed aircraft, allows solving the problem of placing large jet engines under the wing consoles.

The jet engines of the inventive aircraft are connected (see Fig. 17) by the fuel supply lines 44 to the fuel system.

A schematic diagram of the fuel system of the inventive family of aircraft is shown in Fig.17. In this diagram, as well as in the diagrams of the fuel system shown in Figs. 39-41, 47-54, for clarity, the center section compartments are placed in the plane of the drawing (one above the other). This order does not coincide with their order of placement in the claimed family aircraft - along the longitudinal axis of the fuselage.

The fuel system of the claimed aircraft is equipped with two groups of fuel tanks located in the wing consoles. In addition, the aircraft is equipped with a central fuel tank located in the center section.

Each group of fuel tanks of the wing consoles of the aircraft contains the first 45, second 46 and third 47 tanks. The first tanks 45 of the wing consoles are located in the root part of the consoles, the second 46 - on the periphery of the wing consoles, and the third fuel 47 tanks - between them in the middle part of the consoles. The central fuel tank is located in the inner volume of the center wing, its side walls are formed by the front 20 and rear 21 side members and side ribs 25. In addition, to ensure the operation of the fuel system, namely the inter-tank movement of the fuel, it is advisable to tightly connect the mentioned stringers 30 of the closed U-shaped profile with the upper part of the skin 31 of the wing console, and its internal cavity to connect with the first 45 and second 46 tanks of the fuel system.

The fuel tanks of the wing consoles are separated from each other and from the drainage compartment 55 by sealed ribs 26 and 27, and the first tanks of the 45 groups are separated from the central fuel tank 71 by the side ribs 25. The third fuel tanks 47 of the wing consoles of the tanks are separated (see Figs. 6, 18 ) on the main 48 and consumable 49 compartments, the last of which are connected to the fuel supply lines 44 with the engine. One of the walls of the consumable compartments 49 is combined with a sealed rib 24 separating the third fuel tanks 47 from the first 45, and the other with the rear spar 21.

The fuel system of the claimed aircraft is made possible to generate fuel first from the central fuel tank 71, then from the first tanks 45, then from the second tanks 46, and lastly from the third tanks 47 groups (see Fig. 19). As indicated above, the internal volumes of the third fuel tanks 47 are divided into main 48 and consumable 49 compartments, from which the fuel is supplied to the engines. In this case, the fuel system of the aircraft, it is advisable to carry out with the provision of the latest generation of fuel from the consumable compartments (see Fig.20). It is most advisable to ensure fuel production in the indicated sequence in accordance with the claimed decision of the aircraft fuel system, given below.

The claimed decision of the aircraft is made with the possibility of its implementation with three levels of the maximum estimated flight range: with the shortest, intermediate and longest flight ranges in their group.

This can be achieved by the flexible use of the internal volume of the center section - the central fuel tank for placing fuel in an airplane.

In this case, the total volume of two groups of tanks of the wing consoles and the volume of the first compartment 36 of the central fuel tank can be selected from the calculation of providing fueling aircraft with the lowest value of flight range. Using the second compartment 37 of the central tank for fuel placement allows the transition from an aircraft with the shortest flight range to an aircraft with an intermediate flight range: the total volume of two groups of tanks and the volume of the first and second compartments of the central fuel tank can be selected based on the provision of fueling the aircraft with an intermediate range value. The additional use of the third compartment 38 of the central fuel tank allows the transition from an aircraft with an intermediate flight range to the aircraft with the longest flight range: the total volume of two groups of tanks and the volume of the central fuel tank are selected based on the calculation of fueling the aircraft with the largest flight range.

It is advisable to place the said “dry” compartments in the first 45 and third 47 tanks of the tank groups on either side of the sealed ribs 24 separating the first and third fuel tanks. While consumable 49 compartments, it is advisable to place at the rear 21 side members of the wing consoles. For one of the walls of the consumable compartments, it is advisable to use the same sealed rib 24, near which a “dry” compartment is located at the front side member 20. With this placement of the consumable compartment, it is ensured that it is sufficiently removed from the potentially hazardous area during emergency situations during the operation of jet engines, and in combination with the placement of the dry compartment near the sealed rib, it reduces the weight of the structure.

The claimed solution of a passenger aircraft is most appropriate to use for the development of a model range of passenger aircraft, including three groups of aircraft, designed to carry from 58 to 68, from 66 to 83 and from 83 to 108 passengers. On Fig shows an approximate layout of the cabin of passenger aircraft, designed to carry 60, 75 and 95 passengers with a step between rows of seats 812.8 mm If you change the step between the rows of seats and a different layout of passages in the same dimensions of passenger cabins can be placed from 58 to 68, from 66 to 83, from 83 to 108 passengers.

For the indicated groups of aircraft according to passenger capacity, it is advisable to choose a rounded cross-sectional shape of the fuselage, while its equivalent diameter, the same for all aircraft of the family, to choose from the range 3.3 ... 3.55 m. Choosing the equivalent fuselage diameter from this range with a round cross-sectional shape the fuselage allows you to place five seats in a row in the passenger compartment (see Fig.22). At the same time, it is advisable to choose the lengthening of the fuselage of airplanes designed to carry 58 to 68 passengers from the range of 6.5 ... 6.7, calculated for the carriage of 66 to 83 passengers from the range of 7.3 ... 7.6, and lengthening the fuselage of airplanes , designed to carry from 83 to 108 passengers, it is advisable to choose from the range of 8.2 ... 8.6.

In addition, the fuselage, high-speed swept wing and the power plant of the aircraft, it is advisable to choose based on the flight of each aircraft of the family with a cruising speed with a Mach number from 0.75 to 0.82.

The claimed technical solution of the aircraft with the specified dimensions of the fuselage and the wing span of a family of aircraft selected from the range of 25 ... 30 m, allows each aircraft group in terms of passenger capacity to present three aircraft models with the shortest, intermediate and largest maximum estimated range of flight in their group. On Fig, 24, 25 presents the relationship "flight range - commercial load of the aircraft" for three groups of passenger aircraft, designed for transportation from 58 to 68 (Fig.23), from 66 to 83 (Fig.24) and from 83 to 108 (FIG. 25) passengers. Each group of aircraft in terms of passenger capacity includes aircraft with three levels of range: the shortest (curve 50 in FIGS. 23-25), the intermediate (curve 51) and the longest (curve 52) of the maximum estimated flight range. The maximum fueling of the fuel system of the aircraft model range allows for each aircraft passenger group to present aircraft with a maximum estimated flight range of 3100 ... 3300 km, 3900 ... 4100 km and 4600 ... 4900 with a payload of 6000, 7500 and 9500 kg.

The claimed solution of the high-speed swept wing is as follows.

The high speed swept wing consists of a center section and consoles. Each of the wing consoles is composed of six parts 61, 62, 63, 64, 65, 66 (see Fig. 26). Parts of each wing console are delimited from each other by sections parallel to the plane of symmetry of aircraft 67. In this case, the console parts are arranged sequentially along the span of the wing consoles and smoothly mate with each other.

Smooth mating of parts of the wing consoles is achieved by smoothly mating the mid-surfaces of the parts, smoothly changing the maximum relative thickness of the profiles along the span of the wing consoles, smoothly changing the twist angle of the sections and smoothly changing the angle of the transverse V wing.

Parts of the wing consoles are formed as a single spatial system based on a non-planar median surface. Moreover, the profiles of the sections of the console parts, starting from the first 61 from the side section through the fifth 65, are equipped with sections with an S-shape of the middle lines of the profiles. The S-shape of the middle lines of the profiles is composed of two half-waves, one of which 68 is made with negative concavity - the middle lines of the profiles are located below the chord, and the other 69 with a positive one (see Figs. 27-31) - the middle lines of the profiles are located above the chord.

The claimed solution of the high-speed swept wing is characterized in that the half-wave 68 with negative concavity of the S-shaped middle line of the profiles precedes the half-wave 69 with positive concavity when moving along the chord of the wing profile from the nose to the tip of the profile. In this case, the profiles of the sixth part 66 of the console are made with positive concavity of the profiles (see Fig. 33).

The first part 61 of each console is adjacent to the fuselage 1 of the aircraft and is limited by the side section HN and section G-G (see Fig. 26).

, максимальное значение относительной отрицательной вогнутости может составлять при этом минус 0,01 … минус 0,015, а положительной вогнутости 0,001…0,002.On the first part 61 of the wing console, two S-shaped half-waves of the middle lines of the profiles are preceded by an additional half-wave 70 with a positive concavity of the middle lines of the profiles. In the side section in the area adjacent to the nose of the wing profile, the middle line of the profile is located above the chord (made with positive concavity): the length of the additional half-wave with positive concavity can be no more than 5 ... 8% of the length of the chord of the profile, and the maximum value is relative concavity can be no more than 0.001. Further along the chord of the side-section profile, the additional half-wave 70 with positive concavity transforms into an S-shape made of two half-waves 68 and 69, and when passing to the end of the profile half-wave 68 with negative curvature of the midline precedes half-wave 69 with positive concavity of the middle line. The transition from the negative half-wave to the positive half-wave can be performed at that

, the maximum value of relative negative concavity may be minus 0.01 ... minus 0.015, and positive concavity 0.001 ... 0.002.

In addition, in the side section, the value of the maximum relative thickness of the profile is advisable to choose from the range of 14 ... 16% (see Fig. 15), the angle of the geometric twist of the section is from the range of 3 ... 4 degrees (see Fig. 16), the angle of the transverse V wing along the leading edge - from the range of 6 ... 8 degrees, and along the trailing edge - from the range of 8 ... 10 degrees (see Fig. 14).

When moving along the first part 61 of the wing console from the side section to the border with the second part 62 of the console, an additional half-wave 70 with positive concavity of the profiles disappears. The transition from the negative half-wave 68 of the S-shaped shape of the middle lines of the profiles to the half-wave of 69 with a positive concavity of the middle lines of the profiles is shifted to the nose of the wing profiles.

When moving along the first part 61 of the wing console from the side section Н-Н to the border with the second part 62 along the section GG, the maximum relative thickness of the wing profiles, the angle of their geometric twist are reduced (see Fig. 15, 16), and the angle of the transverse V wing increases (see Fig.14).

The second part 62 of the wing console is limited by sections G-G and F-F.

When moving along the second part 62 of the wing console from its border with the first part to its border with the third part 63 of the console, the angle of the transverse V wing continues to increase, reaching a maximum at the boundary between the second and third parts of the wing consoles (see Fig. 14), located close to one quarter of the wing span. Moreover, the angle of the transverse V wing at the border with the third part of the surface along the leading edge is advisable to choose from the range of 9.5 ... 10 degrees, and along the trailing edge - from the range of 14 ... 16 degrees.

In addition, when moving along the second part of the wing console from its border with the first part 61 of the console to its border with the third part 63, two S-shaped half-waves of the middle profile lines fill the entire chord of the profiles (see Figs. 28 and 29) , the transition from the negative to the positive half-wave of the middle lines of the profiles continues to mix to the tip of the profiles, the angle of the geometric twist of the profiles and the maximum relative thickness of the profiles continue to decrease (see Fig. 15, 16).

The third part of 63 wing consoles is limited by sections F-F and D-D. When moving along the third part of the wing console from its border with the second part along the surface FF section to its border with the fourth part of the console along the DD section, the maximum relative thickness of the section profiles continues to decrease (see Fig. 15), reaching at the border between the third 63 and the fourth part 64 of its local minimum. Its value is advisable to choose from the range of 11 ... 12%.

The angle of the transverse V wing, starting from the boundary of the second and third parts of the console along the F-F section, begins to decrease (see Fig. 14). The decrease in the angle of the transverse V wing continues with a further transition to the wing tip along the fourth 64, fifth 65 and sixth 66 parts of the aerodynamic bearing surface. In the end sections of the wing consoles, the value of the angle of the transverse V wing should be selected along the leading edge from the range of 4 ... 5.5 degrees, and along the trailing edge - 5 ... 6 degrees. The angle of the geometric twist of the section profiles on the third part 63 of the aerodynamic bearing surface continues to fall (see Fig. 16).

In addition, during the transition along the third part 63 of the cantilever from its boundary with the second part 62 of the surface along the FF section to its border with the fourth part along the DD section, the transition point of the half-wave 68 with a negative concavity of the S-shaped middle lines of the half-wave profiles in the half-wave 69 s positive concavity.

Under the wing console with a gap relative to it along the border between its third and fourth parts, it is possible to place an axisymmetric elongated body, for example, a jet engine in a nacelle. The boundary between the third and fourth parts of the wing consoles, it is advisable to place near one third of the wing span.

The fourth part 64 of the wing console is limited to sections D-D and CC.

When moving along the fourth part of the wing console from its border with the third part along the D-D section to its border with the fifth part along the CC section, the maximum relative thickness of the section profiles increases (see Fig. 15). On the border between the fourth and fifth parts of the console in section CC, its value can be selected from the range of 12 ... 13%.

.The claimed technical solution of the high-speed swept wing is characterized by the presence on the fourth and fifth parts of the aerodynamic bearing surface of an additional half-wave 70 with positive concavity, which precedes, when passing from the nose of the wing profiles along the chord of the profiles, half-wave 68 with a negative curvature of the S-shape of the middle lines of the profiles. The length of the additional half-wave section 70 with a positive concavity of the middle lines of the profiles from the boundary of the third and fourth parts of the consoles along the DD section increases to 10 ... 20% of the chord length to the border between the fourth and fifth parts of the consoles along the CC section. The length of the half-wave 68 with negative concavity of the S-shaped middle lines of the profiles decreases, and the transition of the negative half-wave 68 of the S-shaped middle lines of the profiles to the positive half-wave 69 continues to shift to the nose of the wing profiles and at the border between the fourth and fifth parts in section С -With it is advisable to perform when

.

The fifth part 65 of the wing console is limited by sections CC and BB.

When moving along the fifth part of the wing console from its border with the fourth part along the CC section to its border with the sixth part along the BB section, the half-wavelength 68 with negative concavity of the S-shaped midline of the profiles decreases to the border between the fifth and sixth parts of the console along the cross-section B-B half-wave 68 with negative concavity disappears. In this case, an additional half-wave 70 with positive concavity of the profiles to section B-B merges with the half-wave 69 of the S-shaped shape of the midline of the profiles (see Fig. 32).

The maximum relative thickness of the profiles, starting from the border between the fourth and fifth parts of the console, decreases. With a further transition to the wingtip on the sixth part 66 of the console, the maximum relative thickness of the profiles continues to decrease to values from the range of 9 ... 11% (see Fig. 15).

The sixth part 66 of the console is limited to sections BB and AA.

The sixth part of the supporting console is made with positive concavity of the midline of the profiles (see Fig. 33).

When moving along the fifth and sixth parts of the cantilever from the boundary between the fourth and fifth parts to the wing tip, the angle of the geometric twist decreases from positive to negative values (see Fig. 16). Moreover, in the end sections, the angle of the geometric twist of the sections, it is advisable to choose from the range of 1.5 ... -2.5 °.

The transition from positive angles of geometric twist of the cross sections to negative can be performed on the border between the fifth and sixth parts of the wing console, close to the section BB.

. При этом в бортовом сечении максимальную относительную толщину целесообразно разместить при

, к границе со второй частью консоли сместить положение максимальной относительной высоты к концу профиля до 37…38%, а к границе с третьей частью консоли дополнительно сместить положение максимальной относительной толщины профиля на хорде крыла к относительной координате

. На последующих частях аэродинамической несущей поверхности максимальные относительные толщины профилей целесообразно располагать при

.The maximum relative thicknesses of the profiles should be located at

. In the side section, it is advisable to place the maximum relative thickness at

, to the border with the second part of the console, shift the position of the maximum relative height to the end of the profile to 37 ... 38%, and to the border with the third part of the console, additionally shift the position of the maximum relative thickness of the profile on the wing chord to the relative coordinate

. On the subsequent parts of the aerodynamic bearing surface, the maximum relative thicknesses of the profiles should be located at

.

. При этом максимальная отрицательная относительная вогнутость профилей уменьшается от значений из диапазона

до 0, где

- относительная координата по хорде профиля.It is advisable to place the section with the maximum negative value of the relative concavity of the profiles on the console parts from the first to the fifth at

. In this case, the maximum negative relative concavity of the profiles decreases from the values from the range

to 0, where

- relative coordinate along the profile chord.

смещается к носку профилей, при этом на пятой части несущей поверхности переход отрицательной полуволны в положительную целесообразно размещать при

.As noted above, in the first to fifth parts, the transition from the negative half-wave of the S-shaped middle lines of the profiles to the positive half-wave from the relative coordinate

it is shifted to the nose of the profiles, while on the fifth part of the bearing surface, it is advisable to place the transition of the negative half-wave into the positive at

.

, при этом на первой части консоли крыла максимальное значение положительной вогнутости S-образной формы средних линий профилей целесообразно разместить при

, при переходе к концевым сечениям смещать его положение к носку профиля, размещая положение участка с максимальным значением относительной положительной вогнутости профилей на шестой части - при

. При этом максимальное значение относительной положительной вогнутости S-образной формы средних линий профилей целесообразно увеличивать от

в бортовых сечениях до

на пятой и шестой частях консолей крыла.It is advisable to place the section with the maximum positive value of relative concavity on the parts of the aerodynamic bearing surface from the first to the sixth at

, while on the first part of the wing console, the maximum value of the positive concavity of the S-shaped middle lines of the profiles should be placed at

, when moving to end sections, shift its position to the nose of the profile, placing the position of the section with the maximum value of the relative positive concavity of the profiles on the sixth part - when

. In this case, it is advisable to increase the maximum value of the relative positive concavity of the S-shaped shape of the middle lines of the profiles from

in side sections to

on the fifth and sixth parts of the wing consoles.

, граница между первой 61 и второй 62 частями консолей крыла (сечение G-G) -

, между второй 62 и третьей 63 частями (сечение F-F) -

, граница между третьей 63 и четвертой 64 частями (сечение D-D) -

, между четвертой 64 и пятой 65 частями (сечение С-С)

, граница между пятой 65 и шестой 66 частями (сечение В-В)

, где z - относительная координата сечения вдоль полуразмаха крыла.It is most advisable to form the claimed swept wing, having sections - the boundaries between the parts of the wing consoles, in the following ranges: side section H-N -

, the boundary between the first 61 and second 62 parts of the wing consoles (section GG) -

, between the second 62 and third 63 parts (section FF) -

, the border between the third 63 and fourth 64 parts (section DD) -

, between the fourth 64 and the fifth 65 parts (section CC)

, the boundary between the fifth 65 and sixth 66 parts (section BB)

where z is the relative coordinate of the section along the wing half-span.

, в сочетании с мотогондолами двигателей, размещенными под консолями крыла, на поверхности крыла реализуется близкое к оптимальному характеру обтекание поверхности (см. фиг.34) и близкое к оптимальному распределение давления по поверхности крыла (см. фиг.35). При этом высокое значение аэродинамического качества (см. фиг.36) реализуется в широком диапазоне чисел Маха (М=0,2…0,82), а максимальное значение показателя К_mах*М достигается при высоких числах М=0,75…0,82.Calculations and aerodynamic experiments show that when using the proposed technical solution for a wing with lengthening λ = 9 ... 11, narrowing, not less than

, in combination with engine nacelles located under the wing consoles, a surface wrap close to the optimum nature (see Fig. 34) and a close to optimal pressure distribution over the wing surface (see Fig. 35) is realized on the wing surface. At the same time, a high value of aerodynamic quality (see Fig. 36) is realized in a wide range of Mach numbers (M = 0.2 ... 0.82), and the maximum value of the indicator K _max * M is achieved at high numbers M = 0.75 ... 0 , 82.

The aerodynamic characteristics of the wing with engine nacelles under the wing when using the claimed technical solution remain stable when used as part of the family of aircraft shown in figures 1, 3-5, which differ in the mass and length of the fuselage (see Fig. 37). As experimental aerodynamic studies and recalculation of their results for airplanes show, the aerodynamic quality of the wing varies within no more than 1 ... 4% when switching from airplanes designed to carry 58 to 68 passengers to airplanes designed to carry from 66 to 83 and from 83 up to 108 passengers, which makes it possible to use the inventive high-speed swept wing without changing the geometric parameters of the wing for the model range of aircraft designed for different passenger capacity and maximum design range of flight.

A comparison of the aerodynamic characteristics of the wing with the characteristics of the wing of one of the foreign analogues (see Fig. 38) shows that the cruising speed of the aircraft with the claimed wing exceeds the cruising speed of the foreign counterpart.

The inventive fuel system of the family of aircraft is arranged as follows.

The inventive fuel system is designed to use the design of aircraft with a positive V wing. The wing includes (see FIG. 6) the front 20 and rear 21 side members, skipped along the wing consoles and the center section 18, the wing skin with its upper 31 and lower 40 parts, sealed side ribs 25 that separate the center section from the wing consoles, ribs 22, 24 , 26, 27 and a longitudinal power set of wing consoles. The center section 18, in addition, is provided with longitudinal walls 34 arranged parallel to the side members. The walls of the center section are provided with cutouts 35. As shown in FIG. 10, the center section can be provided with two walls 34. The cutouts 35 of the walls 34 of the center section can be sealed by their covers 39.

The fuel system comprises a central fuel tank 71 located in the center section and two groups of fuel tanks located in the wing consoles. Schematic diagrams of the fuel subsystems located in the right and left wing console are the same; Figs. 17, 39, 40, 41 show a schematic diagram of the fuel subsystem located in the left wing console.

The Central fuel tank 71 (see Fig.6, 41), made in the form of a caisson, is limited to the front 20 and rear 21 side members, side ribs 25, lower 33 and upper 32 panels. The walls of the center section 34 divide it into compartments. If there are two walls 34 in the center section, the central fuel tank is divided into three compartments 36, 37, 38, the first of which 36 is adjacent to the rear spar, the third 38 - to the front, and the second - 37 is located between them.

Each group of fuel tanks located in the wing consoles includes a first fuel tank 45 adjacent to the sealed side rib 25, a second fuel tank 46 located on the periphery of the wing console, and a third fuel tank 47 located between the first and second tank. The internal volume of the third fuel tank 47 of each group is divided into the main 48 and consumable 49 compartments.

The fuel tanks of the wing consoles are limited by the front 20 and rear 21 side members, the lower 40 and upper 31 parts of the wing skin. They are separated from each other by sealed ribs 24, 26.

The rib 26, separating the second tanks of the groups from the main compartments of the third tanks, is equipped with check valves 83, providing the ability to move fuel from the second tanks to the main compartments of the third tanks and to prevent overflow of fuel in the opposite direction.

One of the walls of the consumable compartment 49 is aligned with the rib 24 separating the third fuel tank 47 from the first fuel tank 45, and the other wall of the consumable compartment is aligned with the wall of the rear spar 21 (see Figs. 41, 18).

In addition, in each group of fuel tanks, a fuel supply line 44 is connected to the engine, the input of which is connected to the third fuel tank 47, namely, to its consumable compartment 49. It is advisable to connect the fuel supply line to the left and right wing console with the fuel ringing line to the main tap ringing, which are not shown in the diagrams. The fuel supply lines 44 to the engine are equipped with booster pumps 72.

The fuel system is equipped with fuel transfer lines 74, 76 with transfer pumps 73. The input of the transfer line 74 is connected to the internal cavity of the first tank, the input of the line 76 is connected to the main compartment 48 of the third tank.

In addition, the inner cavity of the central tank (see Fig. 17) is connected to the inputs of two transfer lines 77, the output of one of them being connected to the third tank of one group of tanks, and the output of the other to the third tank of another group of tanks. In addition, the fuel system may be provided with a transfer line 75, which provides for the transfer of fuel from the second tanks of the fuel system to the third tanks.

The outputs of the fuel transfer lines 74, 75, 77 of the fuel from the first and second tanks and the central fuel tank can be connected, as shown in FIG. 41, to the main compartment 48 of the third fuel tank. The outputs of these lines can be connected, as shown in Fig. 40 and directly with the consumable compartments 49 groups of tanks.

In the walls 53 separating the consumable compartments from the main compartments of the third fuel tanks, and in a sealed rib 26 (see Figs. 39, 40, 41) separating the third fuel tanks from the second, cuts 78 are made. The cuts are made in the upper parts of the walls and ribs , possibly closer to the upper skin of the wing consoles (see FIG. 43).

The transverse power set of each console includes a closed U-shaped stringer 30 (see Figs. 7, 8, 17, 39, 40, 41, 42), hermetically connected to the upper part of the wing skin 31. The internal cavity of the U-shaped stringer 30 (see Fig. 43) is in communication with the internal cavities of the first 45 and second 46 fuel tanks. In this case, it is advisable to place the entrance to the internal cavity of the U-shaped stringer in the upper part of the second fuel tank as close as possible to the upper part of the wing skin.

In addition, each group of fuel tanks is equipped with overflow channels 80, providing the possibility of overflow of fuel from the first fuel tanks to the Central fuel tank. The inlets of the overflow channels should be placed in the upper parts of the first fuel tanks 45.

The inputs of the overflow channels connecting the first fuel tanks to the central fuel tank, it is advisable to place above the output of the U-shaped stringers in the first fuel tanks. The outputs of the overflow channels 80 in the Central fuel tank must be performed with the possibility of overflow of fuel in one of the compartments of the Central fuel tank.

When using the inventive fuel system in aircraft of different flight ranges, the overflow channel 80 may be different. When using the fuel system in an airplane with the shortest flight range, it is advisable to place the outlet of the overflow channel 80, for example, as shown in FIG. 45, in the first compartment 36 of the central fuel tank, and cutouts in the wall of the center section 34 separating the first compartment 36 of the central fuel tank from the second compartment 37, can be overlapped by covers 39. When using the fuel system in airplanes with an intermediate and long flight range, the outlet of the overflow channel 80 can be placed in the second compartment 37 of the central fuel tank, as shown in FIG. In this case, when using the fuel system in an airplane with an intermediate cut-out distance in the center section wall 34 separating the second compartment 37 of the central fuel tank from the third compartment 38, they can be covered by covers. In the model of the aircraft with the longest flight range, the cutouts of the walls of the center section 34, as shown in FIG. 45, do not overlap with covers.

In the lower parts of the walls separating the consumable compartments 49 from the main compartments 48 of the third fuel tanks 47, and the ribs 24 that separate the first fuel tanks from the third, it is advisable to place check valves 81 that provide fuel overflow from the first fuel tanks to the main compartments of the third tanks and from the main compartments of the third fuel tanks into consumable compartments and the exclusion of fuel overflow in the opposite direction.

In addition, the overflow channels 80, it is also advisable to provide check valves 82, providing the possibility of overflow of fuel from the first tanks in the Central compartment and the exclusion of overflow of fuel in the opposite direction.

The slots 78 in the walls of the consumable compartment and sealed ribs 26 separating the third fuel tanks from the second, the inputs and outputs of the internal cavities of the U-shaped stringers 30 and the inputs of the overflow channels 80, it is advisable to perform their height relative to each other. At the same time, it is advisable to place the entrances into the internal cavities of the U-shaped stringers above the slots 78 in the rib 26, which separates the second tanks from the third, which in turn is expedient to place above the slots in the walls of the consumable compartment 49, which separate it from the internal volume of the main compartments 48 of the third fuel tanks .

The inventive fuel system operates as follows.

In the manufacture of the fuel system in accordance with the level of the maximum estimated range of the ordered aircraft, the required value of the maximum fueling of the fuel system is selected. In this case, cutouts in one of the walls of the center section can be hermetically sealed with covers. So, the central fuel tank in the above diagrams is equipped with two walls. This makes it possible to form a fuel system designed for three levels of maximum fueling.

The fuel system with the smallest maximum fueling can be implemented when the covers cover the cutouts of the center section wall closest to the rear wing spar. At the same time, the tanks of the wing consoles and the compartment of the central tank adjacent to the rear spar are refueling.

A fuel system with an intermediate value of the maximum fueling can be implemented when the covers cover the cutouts of the center section wall closest to the front wing spar.

At the same time, the tanks of the wing consoles and the compartment of the central tank adjacent to the rear spar and the compartment between the two walls of the center section are refueling.

Other methods may be proposed for refueling the fuel system when used in aircraft models of the minimum and intermediate maximum estimated flight ranges.

The fuel system with the highest value of the maximum fueling is realized in the absence of covers on the cutouts in the walls of the center section. In this case, the tanks of the wing consoles and the entire central fuel tank are refueling.

After refueling the fuel system and turning on all transfer pumps, fuel circulation between the tanks begins (see Fig. 47).

On the fuel transfer lines 74, 77, fuel from the central tank and the first tanks is supplied to the third tanks of the wing consoles. In this case, fuel from the first group tanks and the central tank can, in accordance with the scheme shown in Figs. 39, 41, be supplied to the main compartments of the third tanks, from which fuel can be pumped to the consumable compartment through the transfer line 76. On the pumping lines 74, 76 and 77 of the first group tanks, the main compartments of the third tanks and the central tank, in accordance with the circuit shown in FIG. 40, it can also be pumped directly to the consumable compartment.

Excess fuel from the consumable compartments through the slots 78 in the walls 53 separating the consumable compartments from the main compartments of the third tanks is moved to the main compartments of the third tanks. Further, through the slots 78 in the sealed walls 26 between the third and second tanks, the fuel moves to the second tanks. In the internal cavity of the stringer 30 with a U-shaped profile, the fuel moves to the first fuel tanks, from where its excess can be transported to the central fuel tank 71 via overflow channels 80.

After turning on the engines, part of the fuel from the consumable compartments begins to flow into the engine.

When the fuel level in the first fuel tanks is lower than the entrance to the overflow channels 80 (see Fig. 48) connecting the first fuel tanks to the central fuel tank, flow compensation from the central fuel tank is stopped by overflow from the first fuel tanks and the central fuel tank is empty.

After emptying the central fuel tank, the total fuel level in the fuel tanks of the wing consoles decreases, and when it is lower than the entrance to the internal cavity of the U-shaped stringers (see Fig. 49), its overfilling into the first fuel tanks ceases. The transfer of fuel from the first fuel tanks to the consumable compartments is not compensated by the overflow of fuel from the second fuel tanks to the first fuel tanks. The first fuel tanks are empty.

After emptying the first fuel tanks, fuel production from the second and third fuel tanks begins. Due to the presence of the check valve 83, the fuel level in the second fuel tank does not exceed the fuel level in the main compartment of the third tank, valve 83 provides overflow of excess fuel from the second tank to the main compartment of the third tank. Since the wing is made with a positive transverse angle V, the second fuel tank is first emptied (see Figs. 50, 51). After emptying the second fuel tank (see Fig. 51), fuel continues to be pumped along line 76 to the consumable compartment. After generating fuel from the main compartment of the third tank, fuel production continues from the consumable compartment (see FIG. 52).

Introduction to the fuel system of the fuel transfer line 75 (see FIGS. 40, 41) from the second fuel tank to the third fuel tank does not change the order of fuel generation from the fuel system tanks. In this case, the alignment of the aircraft changes somewhat.

After generating the first fuel tanks (see FIG. 49), fuel is generated from the second fuel tank (see FIG. 53). The check valve 83 prevents the movement of fuel from the main compartment of the third tank into the second tank of the fuel system. Due to the pumping of fuel, the main compartment of the third tank remains full, production is made only from the second tanks of the fuel system.

After generating fuel from the second tanks (see FIG. 54), fuel is generated from the main compartments of the third tanks and consumable compartments (see FIG. 52).

The claimed solutions of the aircraft, high-speed swept wing and fuel system can be made at the enterprises of the aviation industry.

Claims (22)

1. Aircraft made with the possibility of changing the maximum estimated range, the fuselage of which consists of sequentially placed nose, front, center, rear and tail compartments, the front and rear compartments being made with a length that allows the placement of a different number of passengers on board the aircraft, the chassis includes two main and bow supports, its high-speed swept wing consists of a center section and wing consoles, includes a longitudinal and transverse power set and skin with the lower and upper parts and is made with a positive angle of the transverse V, while the front and rear spars are included in the longitudinal power set of the wing, and the ribs are part of the transverse power set, some of which are sealed, in addition, each wing console is equipped with a dry compartment, one of the walls of which is combined with the front side member, while the wing consoles are made with a decrease in the maximum relative thickness of their profiles during the transition from the side section of the console to the tip, moreover, on the half of the consoles adjacent side sections, there are sections with a local minimum of the maximum relative thickness of the profiles, in addition, the aircraft is equipped with a power plant containing two turbofan air-jet engines installed under the wing consoles symmetrically relative to the plane of symmetry of the aircraft, while the engines are connected by fuel supply lines to the fuel system, equipped with two groups of fuel tanks separated from each other by sealed ribs, each of the groups of tanks including the first and second fuel willow tanks located in the wing consoles, while the first tanks of the groups are located in the root of the wing consoles, and the second tanks on their periphery, and the central fuel tank located in the center section, the side walls of which are formed by the front and rear wing spars and side ribs, characterized in that the installation plane of the turbofan aircraft engines are located near one third of the wing span, and when moving along the wing console from the side section to one quarter of the wing span of the sweep the wing of the aircraft is made with an increase in the angle of the transverse V wing, and with a further transition to the wingtip with its decrease, in addition, the profiles from the side section to at least half of the half-span of the wing are made with positive angles of geometric twist of their sections, and the aforementioned a local minimum of the maximum relative thickness of the profiles of the wing consoles is achieved near one third of the wing span, the mentioned groups of fuel system tanks are equipped with third tanks located in the wing consoles between the first and the second tanks of the groups, the third tanks are divided into main and consumable compartments, the last of which are connected to the fuel supply lines, while one of the walls of the consumable compartments is combined with a sealed rib separating the third fuel tanks from the first and the other with the rear spar, in this case, the fuel system is configured to produce fuel first from the central fuel tank, then from the first tanks, then from the second tanks, and lastly from the third tanks of the groups, and the total volume of two grams Upp tanks and the volume of the entire central fuel tank are selected on the basis of providing fuel for the aircraft with the highest value of the maximum estimated flight range, the aforementioned dry compartments in each wing console are located on both sides of the sealed ribs that separate the first fuel tanks from the third, and in the aforementioned longitudinal power the set of each console included a stringer of a closed U-shaped profile, hermetically connected to the upper part of the wing skin, the internal cavity of which is made communicating with the first and watts the other tanks of the fuel system, while these stringers are located for most of their length near areas with the maximum relative thickness of the profiles of the wing consoles of the aircraft. 2. The aircraft according to claim 1, characterized in that the nose and rear compartments of the fuselage are made with a length that provides the ability to transport by plane from 58 to 68, from 66 to 83 and from 83 to 108 passengers. 3. The aircraft according to claim 2, characterized in that the equivalent diameter of the fuselage of the aircraft is selected from the range of 3.3 ... 3.55 m, while the elongation of the fuselage of the aircraft, designed to carry from 58 to 68 passengers, is selected from the range of 6.5 ... 6.7, designed to carry from 66 to 83 passengers - from the range of 7.3 ... 7.6, and the elongation of the fuselage of the aircraft, designed to carry from 83 to 108 passengers, is selected from the range of 8.2 ... 8.6. 4. The aircraft according to claim 3, characterized in that the fuselage and high-speed swept wing of the aircraft are designed to fly at cruising speed with a Mach number from 0.75 to 0.82. 5. The aircraft according to claim 1, characterized in that in the center wing there are two walls equipped with cutouts parallel to the side members and made with the possibility of hermetically overlapping their covers, while the said central fuel tank is divided into three compartments, the first of which is located at the rear spar, the third at the front spar and the second between the walls, while the total volume of two groups of tanks and the volume of the first compartment of the central fuel tank are selected based on the provision of fueling the aircraft with the smallest shim maximum value calculated flight range, and the total volume of the two groups of tanks and the volume of the first and second compartments of the central fuel tank selected from the software calculating refueling aircraft with an intermediate value of the maximum design range. 6. The aircraft according to claim 1, characterized in that the sealed rib separating the first and third tanks of the wing consoles, at least in its tail section, is located within the first quarter of the wing half-span. 7. A high-speed swept wing, consisting of a center wing and consoles, each of which is composed of six parts, delimited from each other by sections parallel to the longitudinal plane of symmetry of the aircraft, placed sequentially along the span of the wing consoles, seamlessly mated with each other, and formed as a single spatial a system based on a non-planar median surface, while the profiles of the sections of the consoles adjacent to the fuselage are made with an S-shape of the middle lines, composed of two half-waves, one of which of them is made with negative concavity and the other with positive, and the section profiles on the periphery of the wing consoles are made with positive concavity of the midlines, in addition, when switching from the side to the wing end section, the profiles are made with a monotonic decrease in the geometric twist angle from positive to negative values and reducing the maximum relative thickness of the profiles to a value of 9 ... 11%, characterized in that the profiles with an S-shaped middle line are placed on the first five, counting from the fuselage, parts to each console, and when moving along the chord of the profile from its tip to the tip, the half-wave with negative concavity of the middle line of the profiles is preceded by a half-wave with positive concavity, and on the sixth part of each console there are profiles with positive concavity of the middle line, and on the first, fourth and fifth parts each of the wing consoles, two S-shaped half-waves of the midline of the profiles are preceded by an additional half-wave with positive concavity of the midline of the profiles, in addition, near the side section the value of the angle of the transverse V wing along the leading edge is selected from the range of 6 ... 8 °, and when switching from the side section to the boundary between the second and third parts of the consoles, the wing is made with increasing angle of the transverse V wing along the leading edge to values from 9 to 12 °, and upon further transition to the wingtip, the wing angle was reduced to a value from the range 5 ... 6 °, near the side section, the value of the angle of the geometric twist was selected from the range +3 ... + 3.8 °, the maximum relative thickness of the profiles selected however, from the range of 14 ... 16%, at the boundary between the third and fourth parts, a local minimum of the maximum relative thickness of the wing profiles is placed, while during the transition along the fourth part of the console from the border with the third part to the border with the fifth part, the profiles are made with an increase in the maximum relative thickness profiles. 8. The high-speed swept wing according to claim 7, characterized in that the transition from sections with positive values of the angle of the geometric twist of the profiles to sections with negative values of the geometric twist of the profiles is made at the boundary between the fifth and sixth parts of the wing consoles. 9. The high speed swept wing according to claim 7, characterized in that the maximum relative thicknesses of the profiles are located at

where

- relative coordinate along the profile chord. 10. The high-speed swept wing according to claim 7, characterized in that the area with the maximum negative concavity of the profiles is located at

and when passing from the side section to the boundary between the fifth and sixth parts of the wing consoles, the maximum negative concavity of the profiles decreases from the values from the range

to 0, where

- relative coordinate along the profile chord. 11. The high-speed swept wing according to claim 7, characterized in that the relative coordinate of the transition from a half-wave with a negative concavity of the S-shaped middle line of the profiles to a half-wave with a positive concavity of the middle line of the profiles during the transition from the side section to the boundary between the fifth and sixth parts wing consoles shifted out of range

in range

where

- relative coordinate along the profile chord. 12. The high-speed swept wing according to claim 7, characterized in that the section of maximum positive concavity of the profiles is located at

, while moving from the side section to the wingtip, the maximum value of the positive concavity of the profiles increases from

before

where

- relative coordinate along the profile chord. 13. The high-speed swept wing according to claim 7, characterized in that the sweep angle along the leading edge on the first, second and third parts of the wing consoles is selected from the range 31 ... 33 °, and on the fourth, fifth and sixth parts of the wing consoles ... 28 °. 14. The high-speed swept wing according to claim 7, characterized in that the trailing edge on the first, second, third and fourth parts of the wing consoles is made without sweep, and the sweep angle of the trailing edge on the fifth and sixth parts of the wing consoles is selected from the range of 20 ... 23 ° . 15. The high-speed swept wing according to claim 7, characterized in that the coordinates of the side sections of the wing are selected from the range

, the coordinates of the border between the first and second parts of the wing consoles from the range -

, second and third parts -

third and fourth parts

, fourth and fifth parts -

fifth and sixth parts -

from the wing span. 16. A fuel system containing fuel tanks located in the center section and wing consoles, made with a positive angle of the transverse V and includes front and rear side members, equipped with wall cutouts, located in the center section parallel to the side members, the lining with its upper and lower parts, the ribs of the wing consoles some of which are sealed, while the fuel system is equipped with a central fuel tank located in the center section, bounded by the front and rear spars and side ribs and divided by the center section into compartments, and two groups of fuel tanks located in the wing consoles, bounded by the front and rear spars, lower and upper skins and sealed from each other by sealed ribs, while each tank group includes the first fuel tank, separated from the central fuel tank by the side rib, a second fuel tank, located on the periphery of the wing consoles, a third fuel tank, located between the first and second fuel tanks and divided into the main and consumable compartments, one of the walls the fuel compartment is combined with the rib separating the third fuel tank from the first fuel tank, and the fuel supply line to the engine, the input of which is connected to the fuel compartment, in addition, the fuel system is equipped with fuel transfer lines with transfer pumps, the inputs of which are connected to the internal cavities of the first tank and the main compartment of the third tank of each group of tanks, and the inner cavity of the Central fuel tank is connected to the inputs of two pumping lines, and the output of one of them is connected to the first, and the output and the other, with the second group of tanks, in addition, the ribs separating the second tanks from the main compartments of the third tanks are equipped with check valves located near the lower part of the wing skin and made possible to move fuel from the second tanks to the main compartments of the third tanks, characterized in that the second wall of the consumable compartment is aligned with the wall of the rear spar, and in the upper parts of the walls separating the consumable compartments from the main compartments of the third fuel tanks, and in the upper parts of the ribs separating the main compartments of the third tanks from the second, slots are made, while in the longitudinal power set of each console there is a closed U-shaped stringer tightly connected to the upper part of the wing skin, the internal cavity of the specified stringer connected to the internal cavities of the first and second fuel tanks, in addition , each group of fuel tanks is equipped with overflow channels, providing the possibility of overflow of fuel from the first fuel tanks to the Central fuel tank, the inputs of which are located in the upper parts of fuel tanks, and the exits are made possible to transfer fuel to one of the compartments of the central fuel tank, while the exits of the said fuel transfer lines from the first tanks and the central fuel tank are connected to the third tanks of each group, and the exits of the fuel transfer lines from the main compartments of the third tanks are connected to consumable compartments of groups of tanks. 17. The fuel system according to clause 16, wherein each group of fuel tanks is equipped with a fuel transfer line from the second tank to the third. 18. The fuel system according to 17, characterized in that the exits of the fuel transfer lines from the first and second tanks and the central fuel tank are connected to the main compartments of the third tanks. 19. The fuel system according to 17, characterized in that the exits of the fuel transfer lines from the first and second tanks and the central fuel tank are connected to the consumable compartments of the third tanks. 20. The fuel system according to clause 16, characterized in that in the lower parts of the walls and ribs separating the main compartments of the third fuel tanks from the consumable compartments and from the first fuel tanks, check valves are arranged to allow the fuel to move from the first fuel tanks to main compartments of third tanks and from main compartments to consumable compartments and exclusions of fuel movement in the opposite direction. 21. The fuel system according to clause 16, characterized in that the said overflow channels are equipped with check valves, providing the possibility of overflow of fuel from the first tanks into the Central fuel tank and the exclusion of overflow of fuel in the opposite direction. 22. The fuel system according to clause 16, characterized in that the cutouts of the walls of the center section are made with the possibility of hermetic overlapping their covers.

Images