RU2517627C1 - Aircraft - Google Patents

Abstract

FIELD: aircraft engineering.

SUBSTANCE: aircraft comprises fuselage with its tail equipped with two flat sites arranged after another ahead of power plant air intake and turned through obtuse angle relative to each other. Said flats are connected with fuselage skin at an angle to each other without smooth transition. Width of second flat site ahead of said air intake is larger than that of air intake edge. Said second flat site extends on both sides of engine nacelle outer sidewalls beyond air intake edge.

EFFECT: decreased irregularity of supersonic airflow in air intakes.

4 cl, 10 dwg

Description

The claimed technical solution relates to aircraft, mainly to administrative supersonic civilian aircraft designed for business trips, as well as for emergency delivery of small cargoes in order to save time compared to using other vehicles. One of the technical problems to be solved when designing aircraft of this type is to reduce fuel consumption, which can be achieved by reducing the aerodynamic drag of the aircraft and optimizing the operation of the power plant by reducing the degree of unevenness of the supersonic flow supplied to the air intakes.

There are several technical solutions for supersonic aircraft designed for civilian purposes, including the fuselage, wing and power plant located under the wing (see, for example, application for European patent EP 0221204, US applications 20070252028, 20070262207).

The technical solution of a supersonic aircraft (see application for European patent 221204, IPC В64С 30/00, publ. 13.05.1987) contains a fuselage, the bow of which is located in front of the wing, the central part is structurally combined with the wing, the tail of the fuselage protrudes beyond the rear edge of the wing . The nose of the fuselage and partly its central part have side walls inclined inward, forming a surface with a single curvature in the longitudinal direction. The lower surface of the central part of the fuselage is articulated with the lower surface of the wing. Two cylinder-shaped nacelles mounted on the lower surface of the wing on both sides of the fuselage have air intakes located behind the front edge of the wing.

The engine nacelles spaced apart in the wing span in this technical solution partially unload the wing, however, they increase the wave drag and the friction drag of the nacelles by about 20%. This is due to the shape of the nacelles themselves, the mid-section area of which is approximately 1.5 times larger than the entrance area to the air intake, and also to the increase in their wetted surface compared to the layout of engines in a single integrated nacelle. In addition, the use of spaced engine nacelles complicates the task of balancing the aircraft in the event of failure of one of the engines.

The technical solution of a supersonic aircraft in accordance with US applications 20070252028 (IPC V64C 30/00, publ. 11/11/2007) and 20070262207 (B64C 3/50, publ. 11/15/2007) also provides for underwing engine nacelles, but due to the displacement of engines to the aircraft fuselage somewhat simplifies the solution of the problem of balancing the aircraft, and the aerodynamic drag of the aircraft also decreases.

In order to reduce the unevenness of the supersonic air flow at the inlet of the engine’s air intakes and to reduce the influence of the boundary layer accumulated in the stream washing the lower surface of the wing, the technical solutions of this group provide for the placement of engine nacelles on small underwing pylons (see, for example, the decision on US application 20070262207, IPC В64С 3/50, publ. 15.11.2007). Due to the gap between the engine nacelles and the lower surface of the wing, the accumulated boundary layer is discharged from the cut of the air intakes, which reduces the unevenness of the supersonic flow at the entrance to the air intakes, however, the introduction of pylons into the aircraft structure increases the mass of the aircraft.

A number of technical solutions are known for civilian supersonic aircraft (see, for example, US applications 20050224630, 20070145192, 20110315819, US patent 8083171) with wing-mounted engines. In accordance with these decisions, a supersonic aircraft contains a fuselage with a rounded cross-sectional shape, a wing and a power plant located in the rear of the fuselage, equipped with engine cylinders of elongated cylindrical shape and air intakes. Engine nacelles are placed on small pylons on the sides of the rear of the fuselage, engine air intakes are located above the upper skin of the wing consoles. Due to the increase in the area of the surface being washed, the aerodynamic drag increases, and the use of pylons increases the mass of the apparatus.

A known solution of a supersonic aircraft (see application US 2011/0133021, IPC B64C 30/00, B64C 1/00, filing date 09/29/2010, published 09/06/2011) containing a fuselage with a rounded cross section, a wing connected to the rear of the fuselage , and a power plant equipped with engine nacelles and air intakes. The power plant contains two engines combined in a single package and placed in a common engine nacelle. The wing in this solution is placed almost above the fuselage, and the power plant is placed with a gap above the wing in the rear of the fuselage, while the part of the wing located in front of the air intake of the aircraft forms a flat platform. The air intakes of the engines are made in a shape close to the shape of a rectangle. This scheme of the aircraft is characterized by the non-uniformity of the supersonic flow at the inlet to the air intake of the aircraft’s power plant, since during the flow of the supersonic stream, a boundary layer that is not large enough to diverge from the air intake accumulates over a large flat area of the upper wing skin. The location of the air intakes close enough to the influx of the wing makes it possible for vortices to enter the air intakes during take-off and landing modes, which, in turn, also leads to an increase in the uneven flow at the engine inlet. In addition, the presence of a gap between the wing and the engine nacelle increases aerodynamic drag, and the design of the pylon increases the mass of the aircraft.

The closest analogue of the claimed decision of the aircraft is a technical solution known from RF patent 2212360 (IPC ⁷ V64C 30/00, V64C 1/00, application 2002107134/28, filing date 03/21/2002, publ. 09/20/2003). In accordance with this decision, the aircraft contains a fuselage and a power plant located in the rear of the fuselage. The nasal and central parts of the fuselage are made in the cross section of a rounded shape. The power plant is equipped with air intakes, the entrance of which is made with a shape close to the shape of a rectangle. The tail of the fuselage, in addition, is equipped with two flat platforms placed sequentially one after another in front of the air intake of the power plant. Flat platforms are turned relative to each other at an obtuse angle, forming a recess in the rear of the fuselage. The edge of an obtuse angle is oriented perpendicular to the plane of symmetry of the aircraft.

Placing the power plant of the aircraft in the recess of the rear part formed by flat platforms, providing partial shielding of the power plant with the nose and central parts of the fuselage, reduces the aerodynamic drag of the aircraft. However, a drawback of this technical solution is the significant non-uniformity of the supersonic flow at the inlet of the aircraft’s air intake, since the presence of a rarefaction zone at the junction of the rounded fuselage section with the first flat area and a sealing zone at the junction of flat areas leads to significant turbulence of the boundary layer, and the flow moves along flat sites is accompanied by an increase in the boundary layer.

The technical problem solved by the proposed technical solution is to reduce the fuel consumption of the power plant by reducing the degree of unevenness of the supersonic flow at the inlet to the air intake of the aircraft power plant.

The technical problem is solved as follows.

A known aircraft, the fuselage of which has a rounded cross-section, at least in the bow and central parts of the fuselage. The aircraft contains a power plant located in the rear of the fuselage, equipped with air intakes. In addition, the tail of the fuselage is equipped with two flat platforms placed sequentially one after another in front of the air intake of the power plant. In the known solution, flat areas are deployed relative to each other at an obtuse angle.

In the claimed solution, the new one is that the first flat areas and the skin of the rear fuselage are connected to each other without a smooth transition between them, the angle between the first flat area and the second flat area is selected from a range of 170 to 178 degrees, and the width of the second flat area before the cut of the air intakes, the width of the air intakes was chosen to be greater than the width of the air intakes, and the second flat platform was extended on both sides of the air inlet for its cut.

The connection of the first flat platform of the rear of the fuselage with the casing at an angle to each other without a smooth transition between them ensures that there is a sharp edge at the junction of the platform with the casing, which causes a supersonic vortex flow to form along the junction of the platform with the casing, which ensures the suction of the boundary layer accumulated on a flat platform, and draining it away from the fuselage and the air intake. The choice of the width of the second flat area in front of the air intake cut, exceeding the width of the air intake cut, and the extension of the second flat platform on both sides of the air intake for its cut additionally ensures that the boundary layer drains from the zone immediately before the air intake cut.

In addition, the choice of the angle between the first flat area and the second flat area from a range of 170 to 178 degrees also contributes to the formation of a uniform supersonic flow at the entrance to the air intake. When choosing the angle between the first flat platform and the second flat platform, less than 170 degrees, the bottom resistance on the first platform increases, the intensity of the shock wave increases when the flow turns from the first platform to the second, which leads to the appearance of significant areas of flow separation in this angle , which ultimately leads to a decrease in the efficiency of the air intakes. When choosing the angle between the first flat platform and the direction of air supply to the air intake, greater than 178 degrees, the aerodynamic resistance increases over the entire system of platforms, decreases the angle of the shock wave, formed when the flow turns from the first platform to the second, as a result of which it can enter the air intakes. In addition, the Mach number increases beyond this jump, which reduces the efficiency of the air intakes.

The technical result from the use of these techniques is the ability to ensure the formation of the required flow rate at the entrance to the air intakes with a uniform distribution of the velocity field while minimizing aerodynamic drag, which allows to reduce / increase the fuel efficiency of the aircraft.

In addition, it is advisable to choose the length of the second flat platform along the flight direction that is 2.8 ... 5 times higher than the height of the air intakes of the power plant. When choosing the ratio of the length of the second flat platform to the height of the air intakes less than 2.8, the shock wave formed when the flow turns from the first platform to the second can enter the air intakes, which will lead to disruption of the flow structure in the air intakes. To avoid this, it is necessary to reduce the height of the air intakes and increase their width, which will lead to a deterioration in the characteristics of the power plant. When choosing the value of the ratio of the length of the second flat platform to the height of the air intake greater than 5, the thickness of the boundary layer falling at the inlet of the air intake increases, which leads to a deterioration in the characteristics of the air intakes.

In addition, it is advisable to choose the width of the second flat area in front of the air intakes exceeding the width of the air intake adjacent to the side wall of the nacelle by an amount selected from the range from 10 to 20 percent. When choosing the width of the second flat area immediately before the cut-off of the smaller air intakes, the vortex air flow formed along the sharp edge of the junction of the platforms and the fuselage skin will enter the air intake, which will cause deterioration of the air intake characteristics. The choice of the width size of the second flat plate immediately before the cut of the air intakes, of a larger size, will lead to an excessive increase in the washed surface and the midsection of the fuselage, which will lead to an increase in resistance.

The connection of the second flat area with the skin of the rear fuselage at an angle without a smooth transition with the formation of a sharp edge additionally contributes to the outflow of the boundary layer from the second area.

The claimed technical solution is illustrated by the following materials:

figure 1 - General view of the aircraft in isometry;

figure 2 is a view of the rear fuselage in plan;

figure 3 is a side view of the fuselage tail;

figure 4-figure 8 is a cross-section of the fuselage of figure 3;

Fig.9 is an enlarged view of a second flat platform in the place of its extension for the cut of the air intake (view I from figure 2);

figure 10 is a calculated picture of the flow around the rear of the fuselage by supersonic flow.

In accordance with the claimed decision, the aircraft contains a fuselage associated with the wing 1 with a front influx 15, the front horizontal tail 2, the tail vertical tail 3, the power plant. It is most advisable to carry out an aircraft in accordance with the claimed decision according to the aerodynamic scheme "duck".

The sealed fuselage compartments with the cockpit and the passenger cabin are expediently located in the bow and central parts of the fuselage, and to satisfy the requirements for sound impact, wing 1 is expediently connected to the rear of the fuselage. The plumage of the aircraft (see figure 1) consists of a front horizontal tail 2 located in the nose of the fuselage, and a vertical tail 3 located in the rear of the fuselage on the nacelle 4 of the power plant.

The power plant of the aircraft includes engines located in the engine nacelle 4, and air intakes 5. The power plant of the aircraft may contain two, three or more engines. The engines are located in the nacelle in the rear of the fuselage. Engines and air intakes 5 are combined together in a “package”, moreover, as shown in FIG. 1, it is advisable to install the air intakes 5 on top of the rear of the fuselage. Slice 6 of the air intakes can be made in a shape close to the shape of a rectangle, as shown in Figs. 1, 8. This arrangement of the engines makes it possible to minimize the drag of the aircraft, maintain the load-bearing properties of the wing both in flight and in take-off and landing modes, and also reduce balancing losses in the event of a single engine failure.

The nasal and central parts of the fuselage are made in the cross section of a rounded shape (see figure 4).

Part of the fuselage placed in front of the air intakes 5 of the power plant is equipped with two flat platforms 7 and 8, placed sequentially one after the other in front of the air intakes of the power plant. These sites are deployed relative to each other at an obtuse angle Ω (see Fig. 3). The value of the angle Ω - the angle of rotation of the sites relative to each other, in accordance with the claimed solution, is selected from a range from 170 to 178 degrees. The second 8 of these sites, located directly in front of the air intakes of the power plant, is most appropriate to orient in the direction of air supply to the air intake in cruising flight mode.

For a wide class of aircraft intended for flights with supersonic speeds, the angle β (see Fig. 1) between the direction of air supply to the air intakes of the power plant in cruise flight mode and the building horizontal plane 9 of the fuselage lies in the range from 0 to 5 degrees. In the case of using a cylindrical or cylindrical shape, at least in the central part of the fuselage, the direction of the horizontal construction of the fuselage, as a rule, is chosen parallel to the axis of the cylindrical shape. When performing the surface of the fuselage of the aircraft in the form of an elongated body of revolution, the direction of the horizontal construction of the fuselage is chosen parallel to the axis of rotation of the forming surface. In the case of using curved along the horizontal forms of the fuselage of the aircraft, it is advisable to choose the direction of the horizontal construction of the fuselage perpendicular to the plane of the frame located in the flight direction before connecting the first platform to the fuselage skin.

The first 7 of these sites, located at an angle α (see figure 1) to the second site, forms an oblique slice of the fuselage. Given the above range of the angle Ω, the value of the angle α lies in the range from 2 to 10 degrees.

The presence of an oblique cut on the rear of the fuselage formed by the first 7 flat area and the second flat area 8, determines the cross-section of the rear part of the fuselage in accordance with the claimed solution in the form of a combination of rounded and flat parts, and in accordance with the claimed solution, the flat areas of the fuselage are connected to casing at an angle to each other without a smooth transition along the sharp edge 10, as shown in Fig.5-10.

The second 8 flat platform of the rear of the fuselage in accordance with the claimed decision is extended beyond the slice 6 of the air intake of the power plant with the coverage of the side parts of the air intake, as shown in figures 1, 2, 8, 9.

The length (d) of the second flat platform along the flight direction in accordance with the claimed decision, it is advisable to choose 3 ... 5 times higher than the height of the air intake of the power plant (a) (see figure 1).

It is advisable to choose the width of the second flat platform in front of the air intake exceeding the width of the air intake adjacent to the side wall 11 of the nacelle by an amount selected from the range from 10 to 20 percent (see Figs. 8, 9):

2c = (10% ... 20%) × b.

When a supersonic stream flows around an aircraft in the zone of the junction of the first flat area with the fuselage skin, the speed of the supersonic stream increases. At the top of the first flat area, rarefaction waves 13 are formed (see FIG. 10). In the junction zone of the first and second flat areas, the supersonic flow velocity increases and a shock wave 14 is formed. Moving the flow along flat areas is accompanied by an increase in the boundary layer. The angular joint of the flat areas with the fuselage skin along the sharp edge 10 without smooth conjugation forms a vortex flow 12 (see Fig. 2), the boundary layer flows down from the flat areas and is led away from the fuselage and the entrance to the air intake. The boundary layer accumulated in the flow to the area directly adjacent to the air intakes flows down the peripheral parts of the second flat area along the outer side walls of the air intake and is diverted from the fuselage.

The aircraft, made according to the proposed scheme, has higher fuel efficiency characteristics.

Claims (4)

1. Aircraft containing a fuselage, made at least in the bow and central parts with a rounded cross-sectional shape, and located in the rear part of the fuselage power plant, equipped with air intakes, while the fuselage is equipped with two flat platforms placed sequentially one after another in front of the air intakes of the power plant and turned relative to each other at an obtuse angle, characterized in that the first flat platform of the fuselage and the skin of the tail are connected to each other under scrap without a smooth transition, the fuselage platforms are rotated relative to each other at an angle whose value is selected from a range from 170 to 178 degrees, the width of the second flat platform in front of the air intakes is selected to exceed the cut-off width of the air intakes, while the second flat platform is extended on both sides of the outer side walls of the nacelle for cut air intakes. 2. The aircraft according to claim 1, characterized in that the length of the second flat area along the flight direction exceeds the height of the air intakes of the power plant by 3 ... 5 times. 3. The aircraft according to claim 1, characterized in that the width of the second flat area in front of the air intake is selected to exceed the width of the air intake adjacent to the side wall of the nacelle by an amount selected from the range from 10 to 20 percent. 4. The aircraft according to claim 1, characterized in that the second flat area and the skin of the rear fuselage are connected to each other at an angle without a smooth transition.

Images