RU136012U1 - SUPERSONIC PLANE - Google Patents

Abstract

1. A supersonic aircraft containing a fuselage made with a rounded cross-sectional shape and which, when moving against the direction of flight, at least in the bow and tail, is made with a smooth change in the vertical and horizontal overall dimensions of the cross-section, a wing with a front influx located above a fuselage tail end propulsion system equipped with a nacelle with turbojet engines located in it and two supersonic air intakes with a rectangular transverse shape echeniya, wherein the pylon is provided with a plane placed near the plane of symmetry between the aircraft engine nacelle and the rear fuselage skin, thus compressing the air inlets wedges placed vertically near the plane of symmetry plane of the pylon without clearance relative druga.2. A supersonic aircraft according to claim 1, characterized in that the distance from the junction of the leading edge of the influx of the wing with the fuselage to the end of the compression wedges of the air intakes is in the range from 25 to 40 percent of the total length of the fuselage. 3. The supersonic aircraft according to claim 2, characterized in that when moving against the direction of flight from the fuselage section located at the junction of the leading edge of the wing influx with the fuselage, the aircraft fuselage is made with a gradual decrease in the vertical overall dimension of the cross section and a smooth increase in the horizontal overall dimension of the cross section .four. A supersonic aircraft according to claim 3, characterized in that when moving against the direction of flight, starting from a cross section coinciding with the junction of the leading edge of the influx

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.

A number of technical solutions are known for supersonic aircraft (see, for example, RF patents No. 2297371, 2100253, US patent 6857599) with underwing arrangement of the power plant in whole or in part. The solution of a supersonic aircraft according to RF patent 2297371 (IPC B64D 33/02, application form 17.01.2002, publ. 04.27.2007) contains a fuselage with a rounded cross-sectional shape, a wing with a front influx, a power plant equipped with engines and air intakes. The aircraft fuselage is made with a smooth change in the transition against the direction of flight of the vertical and horizontal dimensions of the cross section. Rectangular air intakes of the power plant are located under the wing of the aircraft. This and other solutions with the underwing arrangement of the air intakes of the power plant are characterized by an increased level of noise on the ground.

Noise reduction of jet engines on the ground can be achieved due to the upper placement of the power plant relative to the wing and / or fuselage, which allows to reduce noise on the ground due to the shielding of the engine fan and jet stream directed to the ground.

A number of solutions are known for supersonic aircraft (see, for example, US patents 8083171, 6921045, 6729577, 6698684) with the arrangement of engines and air intakes located in two cylindrical motor-driven aircraft above the wing on the side of the fuselage. The air intakes of the power plant are made axisymmetric in shape with a central body. The engine nacelles in these solutions are placed on pylons mounted on the fuselage in the rear of the aircraft. The disadvantage of these solutions of supersonic aircraft is the significant aerodynamic drag of the aircraft, due to the large surface areas of the pylons and nacelles. In addition, axisymmetric air intakes are characterized by significant sensitivity to the direction of the incoming flow: if the flow runs in a direction different from the axis of the axisymmetric air intake, the characteristics of the air intake are significantly degraded and its efficiency decreases.

Reducing the aerodynamic drag of supersonic aircraft can be achieved by using flat rectangular air intakes in the power plant and placing air intakes in the area of the rear part, partially shielded by the nose or central part of the fuselage.

In accordance with the decision of a supersonic aircraft according to RF patent 2212360 (IPC ⁷ B64C 30/00, B64C 1/00, application 2002107134/28, filing date 03/21/2002, publ. 09/20/2003) the aircraft contains a fuselage and a power located in the rear fuselage installation. The nasal and central parts of the fuselage are made in the cross section of a rounded shape. The power plant is equipped with two air intakes, the inputs of which are made with a shape close to the shape of a rectangle, and are equipped with braking wedges located in the upper part of the air intakes. The tail of the fuselage, in addition, is equipped with two flat platforms placed sequentially one after the other in front of the air intake of the power plant. Flat platforms are deployed relative to each other at an obtuse angle, forming a recess in the rear of the fuselage. Two air intakes are combined in a package in a common engine nacelle, while the braking wedges are located horizontally above the second flat platform, and the air intake shells are a continuation of this horizontal platform. Placing the power plant of the aircraft in the recess of the tail 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, the disadvantage of this technical solution is the significant non-uniformity of the supersonic flow at the inlet to the air intake of the aircraft, since the shock waves from the compression wedges of the air inlets fall on the surface of the fuselage and lead to the separation of the boundary layer. The presence of separation of the boundary layer in the air intakes adversely affects its characteristics and the stability of the engine. In addition, in such an arrangement of the air intakes, the problem of eliminating the mutual influence of the air intakes when one of them is operating in an off-design mode, for example, when the engine is stopped or an air intake surge is encountered, is not solved.

The closest analogue of the proposed solution to a supersonic aircraft is the solution of a supersonic aircraft, known from application for US invention No. 2011/0133021 (IPC B62C 30/00, US NCI 244/13, declared. 09/29/2010, publ. 9.06.2011). This solution of a supersonic aircraft contains a fuselage with a round cross-section, a wing equipped with a front influx and connected to the tail of the fuselage, and a power plant equipped with engine nacelles with turbojet engines and supersonic air intakes placed in them.

When moving against the direction of flight, the aircraft fuselage in this solution, at least in the bow and tail, is made with a smooth change in the vertical and horizontal overall dimensions of the cross section: with an increase in the horizontal and vertical overall dimensions of the cross section of the fuselage in the nose of the aircraft and their subsequent reduction in the middle and rear parts of the fuselage. The fuselage tip in this solution is made in the form of a pointed cone. The cross section of the fuselage, located at the junction of the nose and middle parts of the fuselage, is characterized by the largest overall dimensions of the cross section.

The wing in this solution is placed almost above the fuselage, and the part of the wing located in front of the air intake of the aircraft forms a flat platform. The supersonic air intakes of the power plant are made in a shape close to the shape of a rectangle and are located in the area of the front influx of the wing. The compression wedges are located horizontally at the bottom of the air intakes. In the solution under consideration, between the parts of the lower skin of the nacelle adjacent to its side walls and the wing surface, drain slots of the boundary layer of small height are placed. In the solution under consideration, the air intake entrances are located quite close to the front wing influx, so the distance from the junction of the leading edge of the influx with the fuselage to the ends of the compression wedges is from 5 to 12 percent of the total length of the fuselage.

As in the analogue discussed above, the fuselage in the middle and rear parts partially screens the power plant. The horizontal placement of the compression wedges of the supersonic air intake in its lower part excludes the occurrence of shock waves from the compression wedges in the air intake, thereby reducing the degree of non-uniformity of the supersonic flow at the entrance to the air intake of the aircraft. However, the location of supersonic air intakes close enough to the influx of the wing leads to the possibility of vortices falling from the influx into the air intakes during takeoff and landing modes. The presence of extended wing surfaces and fuselage skin in front of the air intakes leads to the accumulation of a significant thickness of the boundary layer before entering the air intakes, which, combined with the insufficient height of the slots of the boundary layer drain, leads to the inhibition of the flow to the compression wedges, which leads to disruption of their operation in the design mode . These factors lead to a significant uneven flow in the air intakes. In addition, this scheme of a supersonic aircraft does not solve the problem of eliminating the mutual influence of the air intakes when one of them operates in off-design mode, for example, when the engine is stopped or when a supersonic air intake surges.

The technical problem solved by the proposed technical solution is to eliminate the interference of supersonic air intakes on each other in off-design flight modes in combination with reduced fuel consumption.

The technical problem is solved as follows.

The well-known supersonic aircraft contains a fuselage with a rounded cross-sectional shape, a wing with a front influx, and a power plant located above the tail of the fuselage, equipped with a nacelle with turbojet engines and supersonic air intakes located in it. In the known solution of a supersonic aircraft, when moving against the direction of flight, the aircraft fuselage, at least in the bow and tail, is made with a smooth change in the vertical and horizontal overall dimensions of the cross section. The air intakes of the power plant are made with a rectangular cross-sectional shape. In the known solution, the new one is that the aircraft is equipped with a pylon located near the plane of symmetry of the aircraft between the nacelle and the skin of the rear fuselage. Wedges of compression of the air intakes are placed vertically near the plane of symmetry of the aircraft above the pylon without a gap relative to each other.

The vertical placement of the compression inlets of the air intakes near the plane of symmetry of the aircraft without a gap relative to each other above the pylon provides almost complete elimination of the mutual influence of the air intakes on each other, since the shock waves from the compression wedges extend along the upper surface of the wing outward from the nacelle, which not only eliminates their influence on each other on off-design flight modes, but additionally reduces aerodynamic drag. The pylon placement of the power plant eliminates the accumulation of the boundary layer in supersonic air intakes, which, along with a decrease in the supersonic flow unevenness at the inlet of the air intakes, also reduces the required air flow through the boundary layer suction system from the compression wedges compared to the lower location of the compression wedges, which also reduces aerodynamic drag

The technical result from the use of the proposed methods is the elimination of the mutual influence of supersonic air intakes on each other in off-design flight modes in combination with a decrease in fuel consumption due to a decrease in aerodynamic drag and the degree of unevenness of the air flow at the entrance to supersonic air intakes.

In addition, in the claimed solution, the distance from the junction of the leading edge of the influx of the wing with the fuselage to the end of the compression wedges of the air intakes can be selected from a range of 25 to 40 percent of the total length of the fuselage. Such a mutual arrangement of the influx of the wing and the entrance to the air intake of the power plant significantly reduces the effect of vortices from the influx on the air flow supplied to the air intakes during takeoff and landing modes, which further reduces the uneven flow at the engine inlet.

In addition, when moving against the direction of flight from the fuselage section located at the junction of the leading edge of the wing influx with the fuselage, the aircraft fuselage can be made with a smooth decrease in the vertical overall dimension of the cross section and a smooth increase in the horizontal overall dimension of the cross section, which provides shielding of the power plant with the fuselage and noise reduction on the ground.

Most preferably, the reduction in the vertical overall dimension of the fuselage cross sections when moving against the direction of flight, starting from the cross section coinciding with the junction of the leading edge of the wing influx with the fuselage, to the cross section coinciding with the ends of the braking wedges, is selected from a range of 25 up to 35 percent, and up to a cross section that coincides with the ends of the side shells of the air intakes, by a value comprised in the range from 40 to 50 percent.

Most preferably, the magnitude of the increase in the horizontal overall size of the fuselage when moving against the direction of flight, starting from the cross section coinciding with the junction of the leading edge of the wing influx with the fuselage, to the cross section coinciding with the ends of the braking wedges, is selected from a range of 15 to 25 percent, and to a cross section that coincides with the ends of the side shells of the air intakes, from the range enclosed in the range from 20 to 30 percent.

The claimed technical solution is illustrated by the following materials:

figure 1 - General view of the aircraft in isometry;

figure 2 is a General view of the aircraft from the side,

figure 3 - General view of the aircraft in terms of plan,

figure 4 - General view of the aircraft against the direction of flight,

figure 5-figure 9 - cross sections and sections of the fuselage of figure 2,

figure 10 is a General view of the rear of the fuselage in isometry.

In accordance with the claimed decision, the aircraft contains a fuselage 1, paired with a wing 2 with a front influx 3, tail 4, power plant and pylon 8.

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 2 is expediently connected to the rear of the fuselage.

The plumage 4 of the aircraft can be performed according to the T-shaped pattern, as shown in figures 1 and 10, although other layouts of aerodynamic controls can be used in the considered solution.

Pylon 8 is located near the plane of symmetry 7 of the aircraft between the nacelle and the skin of the rear fuselage.

The power plant of the aircraft is located above the rear of the fuselage. The power plant contains a nacelle 5 with turbojet engines located in it and two supersonic air intakes with a rectangular cross-sectional shape. Wedges of compression 6 of the air intakes are placed vertically near the plane of symmetry of the aircraft 7 above the pylon 8 without a gap relative to each other (see Fig.8-10).

Most preferably, the joint of the leading edge of the influx of the wing with the fuselage is marked in the cross section of the fuselage with a relative coordinate:

,

,

where Lx is the coordinate measured from the nose of the fuselage, Lf is the total length of the fuselage, the value of which is in the range from 50 to 60 percent. In this case, the nacelle 5 is most preferably displaced in the direction opposite to the flight relative to the joint 9 of the leading edge of the influx of the wing with the fuselage, choosing the distance from the joint of the leading edge of the influx of the wing with the fuselage to the end of the compression wedges of the air intakes from a range of 25 to 40 percent of total fuselage length.

The fuselage 1 of the aircraft is made with a rounded cross-sectional shape. When moving against the direction of flight, the fuselage of the aircraft is made with a smooth change in the vertical and horizontal overall dimensions of the cross section, as shown in Fig.4-9. In this case, when moving against the direction of flight from the fuselage section located at the junction 9 of the leading edge of the wing influx with the fuselage, the aircraft fuselage is made with a gradual decrease in the vertical overall dimension of the cross section and a smooth increase in the horizontal overall dimension of the cross section.

Most preferably, the vertical overall dimension of the cross section of the fuselage when moving against the direction of flight, starting from the cross section coinciding with the place 9 of the junction of the leading edge of the wing influx with the fuselage, to the cross section coinciding with the ends 10 of the braking wedges, is reduced by 25 to 35 percent , and up to a cross section coinciding with the ends of the 11 side shells of the air intakes, by 40 to 50 percent. The horizontal overall cross-sectional dimension of the fuselage is most preferably increased by 15 to 25 percent and by 20 to 30 percent, respectively.

In flight, when a supersonic flow flows around an airplane on the aircraft structural elements, shock waves, fans of rarefaction waves, vortex flows, and other flow features occur.

In take-off and landing modes and some other flight modes, a vortex flow occurs at the leading edge of the wing influx, which, due to the remote placement of air intakes from the wing influx, is diverted from the air intakes.

When flowing around the upper surface of the fuselage, the accumulated boundary layer is discharged from the air intakes into the space between the nacelle and the fuselage.

When a plane is flying at cruising speed with a supersonic speed around the tail of the aircraft, a supersonic flow structure is formed.

Firstly, during supersonic flow around braking wedges, shock waves form on them. These jumps propagate at a relatively small angle to the plane of symmetry of the aircraft in the opposite direction from the front edge of the vertical wedges of the air intakes. The indicated shock waves close in one jump in the region of the outer shell of the air intakes. These combined jumps extend further above the upper surface of the wing.

Secondly, shock waves occur when a pylon flows around a free stream. This shock wave is located in the area between the fuselage and the engine nacelle and extends at a relatively small angle to the plane of symmetry of the aircraft in the opposite direction to the flight. The presence of this shock wave leads to a slight increase in the thickness of the boundary layer in the region of its influence.

Thirdly, in the region of the trailing edges of the wing and the fuselage there are closing shock waves of the seal. The closing shock of the fuselage seal interacts with the rarefaction region of the flow, which is located on the lower surface of the engine nacelle behind the tail of the fuselage. As a result, pressure in this area is growing somewhat.

Claims (5)

1. A supersonic aircraft containing a fuselage made with a rounded cross-sectional shape and which, when moving against the direction of flight, at least in the bow and tail, is made with a smooth change in the vertical and horizontal overall dimensions of the cross-section, a wing with a front influx located above a fuselage tail end propulsion system equipped with a nacelle with turbojet engines located in it and two supersonic air intakes with a rectangular transverse shape echeniya, wherein the pylon is provided with a plane, disposed near the plane of symmetry between the aircraft engine nacelle and the rear fuselage skin, thus compressing the air inlets wedges placed vertically near the plane of symmetry plane of the pylon without a gap relative to each other. 2. The supersonic aircraft according to claim 1, characterized in that the distance from the junction of the leading edge of the influx of the wing with the fuselage to the end of the compression wedges of the air intakes is in the range from 25 to 40 percent of the total length of the fuselage. 3. The supersonic aircraft according to claim 2, characterized in that when moving against the direction of flight from the fuselage section located at the junction of the leading edge of the influx of the wing with the fuselage, the aircraft fuselage is made with a smooth decrease in the vertical overall dimension of the cross section and a smooth increase in the horizontal overall dimension cross section. 4. A supersonic aircraft according to claim 3, characterized in that, when moving against the direction of flight, starting from a cross section coinciding with the junction of the leading edge of the wing influx with the fuselage, to a cross section coinciding with the ends of the braking wedges, the vertical overall dimension of the cross section the fuselage is reduced by an amount enclosed in the range from 25 to 35 percent, and to a cross section coinciding with the ends of the side shells of the air intakes, by an amount enclosed in the range from 40 to 50 percent. 5. A supersonic aircraft according to claim 3, characterized in that, when moving against the direction of flight, starting from a cross section coinciding with the junction of the leading edge of the wing influx with the fuselage, to a cross section matching the ends of the braking wedges, the horizontal overall dimension of the cross section the fuselage is increased by a value enclosed in the range from 15 to 25 percent, and to a cross section coinciding with the ends of the side shells of the air intakes, by an amount enclosed in the range from 20 to 30 percent.

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