RU2670664C9 - Asymmetrical air-scoop for three flow engine of faster-than-sound aircraft - Google Patents

Abstract

FIELD: machine building; aviation.SUBSTANCE: invention relates to inlets of high-speed aircrafts. Asymmetrical air-scoop for a three flow engine of faster-than-sound aircraft contains a spatial vee (1), feedwell (2), side walls (3), subsonic diffuser (6), neck finish and control system of boundary layer. In subsonic diffuser (6) there is axial parting channel (10) with a leading edge (11) of complex shape for draining low-power part of the flow with large losses to outbound of the air-scoop. To control the flow in the internal channel of air-scoop there is a wedgelike opening of draining (4) located in slicing of air-scoop neck finish. Side walls of air scoop are made with a sweepforward of leading edges, located between compression ramps of vee and feedwell. Feedwell is made with a sweepforward of leading edge.EFFECT: invention reduces weight and increases reliability in a wide range of in-flight parameters.3 cl, 8 dwg

Description

The invention relates to aircraft, in particular, to the designs of input devices of high-speed aircraft.

The problem of creating an effective power plant for an aircraft is inextricably linked with the need to ensure effective braking of the flow in the air intake in a wide speed range, including at supersonic speeds. To do this, it is necessary to ensure the minimum loss of total pressure along the path of the air intake and subsonic diffuser and the maximum uniformity of flow in front of the engine over the entire speed range. It is necessary to reduce the resistance of the air intake and subsonic diffuser during their integration with the aircraft body.

Recently, searches have been made for circuit and design solutions for supersonic passenger airplanes that satisfy the requirements of low sound impact when flying with a Mach number of more than 1 and a low noise level during take-off and landing conditions, provided that operation is economical.

Known subsonic air intake (CN 102923309, 2013), recessed into the fuselage. This arrangement provides minimal drag. In the air intake there is a bypass channel, which carries out the bypass of the boundary layer from the fuselage for the propulsion system. The bypass channel is located under the air intake along its entire length, which negatively affects the value of the total pressure coefficient. In addition, the air pumped through the bypass channel does not participate in the cycle of the propulsion system, which means that the energy is spent on braking the flow, pumping and discharging air without an energy contribution to the operation of the power plant.

Known air intake (US 5749542, 1998) with improved performance at supersonic flight speeds. Before entering the air intake is a spatial rounded protrusion with isentropic surfaces. With this solution, it is possible to improve the performance of the air intake by deflecting the low enthalpy boundary layer from the fuselage to the sides of the entrance. In the outlet cross section of the air intake, uneven distribution of characteristics should inevitably appear. The effective operation of such a device is observed in the design flight regimes in terms of speed and design angles of flow inflow. The boundary layer grows on the walls of the air intake, which will also lead to an increase in the level of unevenness.

Studies on the use of the power plant of a supersonic business aircraft (SDS) of an integral type using variable-cycle engines (DITs) of various schemes, including a motor scheme with an adaptive fan and a third circuit channel, show the expediency of using the latter. In particular, you can consider a circuit with a double-circuit air intake and a three-circuit engine and an adaptive fan.

The presence of several circuits at the engine inlet allows you to significantly change the reduced flow rate at the engine inlet, depending on the operating mode. The specific use of this feature of the engine depends on the general concept of the power plant (selection of the design mode for determining the size of the air intake, etc.). So, the size of the air intake can be selected in cruise mode using only the internal circuit, and the larger required value of the reduced air flow during take-off mode is ensured by opening additional air intake flaps from the external circuit.

A large air flow through the engine provides a small external resistance of the air intake, the dimensions of which are selected according to the total value of the engine flow, while the main circuit operates in cruising mode as a dual-circuit engine with a small bypass ratio.

Double-circuit air intakes are used in engines with an additional third circuit with a change in the cycle of operation. The power plant (CS) scheme with a third circuit is a dual-circuit engine with an adjustable second fan stage and a third circuit. Behind the first stage of the fan there is an adjustable air bypass system in the main or third circuit, depending on the operating mode.

The channel of the third circuit can be used as a channel for transferring low-energy air or a boundary layer without supplying energy to it (supersonic cruising and transonic modes), or as a third circuit of a dual-circuit engine with energy supply (take-off mode).

In supersonic cruising mode, by controlling the compression ratio of the second fan stage, the engine can be switched to a mode with a lower bypass ratio and a higher specific thrust. A lower required air flow rate at the engine inlet can be obtained from the core of the air intake stream in order to reduce unevenness and increase the pressure recovery coefficient at the engine inlet. Low-energy air from the air intake enters the third circuit.

In transonic mode, moving the second stage of the fan to the “on” position provides a low bypass ratio and a high thrust value.

In the take-off low-noise mode, the engine provides low values of the jet velocity. The third circuit jet forms an acoustic screen to reduce the noise level of the jet from the nozzle of the bypass part of the engine.

In a double-circuit air intake, the main circuit in cruising mode operates under conditions of lower losses of full pressure and significantly less uneven flow at the inlet. The engine also has higher fan efficiency values for the main circuit. The third motor circuit is used as a reduced pressure air bypass channel. In the take-off low-noise mode, the field unevenness at the engine inlet is low, both engine circuits operate, providing a given engine thrust at a low jet speed. The above features open up new prospects for the application of the engine circuit with a dual-circuit air intake and DIC.

The most advanced Gulfstream VDS concept in development offers a dual-circuit intake circuit (Conners, T. R., and Howe, D. C., "Supersonic Inlet Shaping for Dramatic Reductions in Drag and Sonic Boom Strength," AIAA Paper 2006-0030, Jan. 2006), which satisfies two basic requirements: low sound impact intensity with good performance in terms of drag and lift in cruising flight mode. This is ensured by profiling the circuit of the power plant, and more precisely the engine nacelle SU, which has small angles of inclination of the outer generatrix, and, therefore, low resistance. To ensure high performance SU and the necessary bypass at the stage of acceleration, in the nacelle there is an internal path for passing bypassed air and ejecting it in the tail of the SU at low flight speeds. In this case, the nacelle is designed in such a way that all the shock waves that form from the compression surfaces remain inside the nacelle, and the front edge of the air intake shell is designed taking into account the requirements for the air intake. The bypass channel turns out to be large, but as a result, the resistance of the engine nacelle (due to small external tilt angles) decreases, the characteristics of the air intake improve and, in general, the characteristics of the control system improve. However, the disadvantage of this concept is the use of the axisymmetric design of the nacelle, which requires the use of an aircraft circuit (A) with the placement of engines on pylons. The latter ultimately increases the total resistance of the aircraft.

The closest analogue selected for the prototype is an air intake with variable geometry for a supersonic aircraft (RU 149896, 2014). The air inlet contains a spatial wedge, a casing, a boundary layer drainage system, an air inlet throat with a subsonic diffuser, where through windows are provided on the side walls, a channel behind the air inlet neck is made in the form of a hexagonal section at the inlet and is formed to exit into a subsonic curved diffuser with a circular cross section, the system drainage of the boundary layer has through transverse cracks and perforation on the wedge. The technical solution provides automatic control of the flow path for expanding the range of flight modes and Mach numbers, improving the characteristics of the air intake, automatically eliminating the occurring air flow breaks in the tract, increases the reliability of the air intake integrated with the aircraft body, and reduces the radar visibility of the air intake. However, a thick vortex layer accumulating on the aircraft fuselage in front of the air intake significantly reduces the average total pressure coefficient and increases the uneven flow in front of the engine.

The technical problem solved by the claimed invention is to expand the arsenal of technical means, namely, the creation of an asymmetric dual-circuit air intake for working with a three-circuit engine.

The technical result provided by the invention is to realize its purpose, i.e. in creating an asymmetric dual-circuit air intake, which reduces the loss of full pressure, reduces the size of unevenness and pressure pulsations in front of the engine fan.

The claimed technical result is achieved due to the fact that the asymmetric air intake contains a spatial wedge, a shell, side walls, a subsonic diffuser, as well as a throat and a boundary layer control system, where the wedge consists of compression panels and the boundary layer control system is equipped with a wedge-shaped slot a drain located in the throat section, at the outlet of the subsonic diffuser there is an axial separation channel, the front edge of which is made of a complex shape, described by the formula:

Where

D is the diameter of the separation channel, m;

- высота выступа, м;

- the height of the protrusion, m;

ϕ _CR - the angle of conjugation of the ledge right, deg;

ϕ _lev - the angle of conjugation of the protrusion of the left, deg;

ϕ _center — angle at the apex of the protrusion, degrees;

- функция Хевисайда.

- Heaviside function.

The side walls of the air intake are made with a reverse sweep of the leading edges and are located between the compression panels of the wedge and the shell, and the front edge of the shell is made with a reverse sweep.

The significance of the distinguishing features of the claimed asymmetric air intake is confirmed by the fact that only the totality of all the design features describing the invention is sufficient to solve the technical problem and achieve the claimed technical result. Namely:

- the boundary layer control system contains a wedge-shaped drainage slot located in the throat section, which allows controlling the flow in the internal channel of the air intake due to the pressure differential in the throat region and in the surrounding stream;

- an axial separation channel is placed at the outlet of the subsonic diffuser, which is actually mated to a three-circuit engine, the size of the internal circuit is selected so that 30% of the total air intake is directed to the external circuit;

- the front edge of the separation channel is made of a complex shape, described by formula (1), for draining the low-energy part of the stream with large losses into the external circuit of the air intake, which allows you to drastically increase the level of total pressure in the Central circuit of the air intake;

- the side walls and the front edge of the shell are reversed sweep, which reduces the mass of the air intake, compactness and improves the characteristics of the air intake at launch and take-off mode, ensures stable operation in cruising mode and the implementation of autostart air intake in all flight modes.

Thus, the technical result provided by each of the essential features of the present invention, in combination, allows you to create an asymmetric double-circuit air intake, which reduces the loss of total pressure, reduces the size of unevenness and pressure pulsations in front of the engine fan.

The present invention is illustrated by the following detailed description of the design of the asymmetric air intake for a three-circuit engine and its operation with reference to the illustrations shown in FIG. 1-5, where

in FIG. 1 shows the design of an asymmetric air intake;

in FIG. 2 is a longitudinal section of FIG. one;

in FIG. 3 is a plan view of FIG. one;

in FIG. 4 is a front view of FIG. one;

in FIG. 5 - shape of the leading edge of the axial separation channel;

in FIG. 6 is a picture of Mach numbers in the longitudinal plane of symmetry of the air intake;

in FIG. 7 is a picture of the distribution of the total pressure coefficient in the outlet section of the air intake;

in FIG. 8 - dependence of the coefficient of total pressure at the outlet of the central circuit on the coefficient of total pressure at the outlet of the outer circuit of the air intake.

An asymmetric double-circuit air intake for a three-circuit engine contains a spatial wedge 1 consisting of compression panels, a shell 2, side walls 3, a boundary layer control system that has a wedge-shaped drain slot 4 located in the throat section 5, a subsonic diffuser 6 (Fig. 1, 2 ) The wedge 1 is made in a projection with a width equal to the width B of the air intake, mounted on the wall 7 of the aircraft body and has front swept edges 8 (Fig. 3). The angle β of the wedge 1 in the vertical plane relative to the longitudinal axis X is 5-15 °.

The front edge 9 of the shell 2 is made with reverse sweep with an angle of sweep χ relative to the transverse axis Z. The inner surface of the shell 2 is made of adjacent planes forming a confuser surface. The side walls 3 are made with reverse sweep of the leading edges, which begin from the wedge 1, where its projection width reaches a value equal to the width B of the air intake channel and extending to the shell 2. An axial separation channel 10 is located in the subsonic diffuser 6, which has a leading edge 11 complex shape (Fig. 4, 5), and defined by the formula:

,

,

Where

D is the diameter of the separation channel, m;

- высота выступа, м;

- the height of the protrusion, m;

ϕ _CR - the angle of conjugation of the ledge right, deg;

ϕ _lev - the angle of conjugation of the protrusion of the left, deg;

ϕ _center — angle at the apex of the protrusion, degrees;

- функция Хевисайда.

- Heaviside function.

The inventive asymmetric dual-circuit air intake for a high-speed aircraft operates as follows. The incident supersonic flow is inhibited in oblique shock waves initiated by the spatial wedge 1, intersecting in the plane of symmetry and interacting with each other, forming a flow close to conical. The resulting total shock wave passes near the leading edge 9 of the shell 2, the shape of which and the sweep angles χ in the plane parallel to the XOZ plane are selected taking into account the configuration of the total shock wave formed on the calculated Mach number of the air intake according to the gas-dynamic cut-out from the current tube for the calculated multi-jump braking circuit.

When braking the flow on the spatial wedge 1, the transverse pressure gradient and the velocity component in the direction of the Z axis act on the boundary layer formed on the aircraft body and the wedge 1 and reduce its thickness by draining to the sides in front of the side walls 3. The flow on the wedge 1 is spatial and therefore the flow around the side walls 3, the inner surface of which is inclined to the plane of symmetry of the air intake, is also realized with the formation of shock waves in the channel of the air intake.

Further inhibition of the supersonic flow occurs in oblique shock waves formed on the confuser surface of the shell 2 and when flowing around the side walls 3 of the air intake, as well as in the closing shock wave, which is close in intensity to the direct shock located near the throat section 5. Due to the discharge of the boundary layer through the slot 4 on the greater part of the length of the spatial wedges 1, the side walls 3 with arrow-shaped edges, the necessary air bypass to the outside is ensured in the launch mode, i.e. it starts up at the estimated Mach number and at flight speeds lower than the calculated ones.

One of the main points in the operation of the air intake is the separation of the flows in the subsonic diffuser using the axial separation channel 10. In this case, the flow with high total pressure enters the central circuit of the air intake and further into the engine, the flow with low total pressure separated by the separation edge 11 enters external circuit of a three-circuit engine.

The invention provides:

- improving the characteristics of the air intake for a supersonic aircraft;

- reduction in mass of the air intake, its compactness;

- automatic regulation of the flow path to expand the range according to flight modes and Mach numbers;

- ensuring the automatic elimination of arising air flow gaps in the subsonic diffuser;

- increasing the uniformity of the air flow in front of the engine;

- improving the reliability of the air intake integrated with the aircraft;

- reducing the spread of noise from the engine compressor.

This scheme has the following advantages:

- the geometrical contour of the air intake is selected from the gas-dynamic cut-out of the current tube during flow braking in the system of oblique spatial shock waves and the closing normal jump in front of the curved subsonic diffuser, which makes it possible to organize flow braking with minimal losses for a wide range of flight Mach numbers, i.e. the streamlines behind the oblique shock wave system intersect with the assumed transverse contour of the air intake. The result is a complex three-dimensional surface, forming the appearance of an air intake;

- initial braking of the oncoming flow is carried out on the compression wedge with an angle of rotation of the flow from 8 ° to 15 ° with respect to the axis of the fuselage in the plane of symmetry of the air intake. Further braking occurs in the shock wave reflected from the shell;

- the subsonic diffuser is made slightly curved with a length

L _DD / D _DD , _out = 2 / 0.7 = 2.86 m,

Where

L _DD - the length of the subsonic diffuser,

D _DD , _o - the diameter of the subsonic diffuser at the output.

The output cross section of the subsonic diffuser along its length increases 1.6 times, and the vertical displacement of the axial line of the subsonic diffuser from the condition of integration with the fuselage body is 0.22 m. The trapezoidal channel of the subsonic diffuser smoothly passes over the section of the throat of a constant area of 0.243 m ² into a circular diameter of 0.85 m. The geometry of the channel contour is approximated by three-dimensional splines of the third order. In the output section of the subsonic diffuser, a cylindrical section about 0.35 m long is docked.

Extensive design studies with the free-stream Mach number equal to the critical Mach number (M _cr ) of multi-mode unregulated air intakes integrated with the aircraft and designed for cruising flights with Mach numbers from 1.6 to 2.0 were carried out using standard certified FLUENT and FASTRAN software packages based on the solution of stationary Navier-Stokes equations averaged by Reynolds for three-dimensional viscous flows. The “k-ε" model was chosen as the turbulence model.

Computational studies of the flow and characteristics of the air intake with a subsonic curved diffuser in the non-throttling modes and with the throttling of the flow in the outlet section, performed using the ESI-CFD software package based on solving the Reynolds averaged Navier-Stokes equations, have shown the efficiency and correctness of the proposed technical solution. The values of the total pressure recovery coefficient σ at flight Mach numbers from 0.6 to 2.0 are 0.96-0.92 with a discharge air flow rate ranging from 1.5% to 2.5% of the total flow rate.

In FIG. 6 shows a picture of Mach numbers in the longitudinal plane of symmetry of the air intake at cruising mode with a Mach number of 1.8, and in FIG. 7 shows the distribution of the total pressure coefficient in the outlet section of the double-circuit air intake. It is seen that a vortex zone is formed at the bottom of the channel with increased total pressure losses, which, due to the presence of a dividing flap, merge into the external contour of the air intake. In FIG. Figure 8 shows the dependences of the values of the coefficient of total pressure at the outlet of the central circuit on the coefficient of total pressure at the outlet of the external circuit for various values of backpressure at the outlet. It can be seen that with optimal throttling of the air intake circuits, it is possible to obtain high values of the total pressure coefficient at the inlet of the central circuit at the level of σ _int from 0.9 to 0.91.

Thus, the proposed dual-circuit asymmetric intake for a supersonic aircraft with a three-circuit engine provides effective braking of the incoming air flow and reducing its resistance with minimal mechanical control of the values of its design parameters and discharge of the boundary layer in operating and launch modes, and a uniform profile of flow parameters in the output cross section of a subsonic diffuser. The air intake meets all the necessary requirements for working with a three-circuit engine and provides high performance in all engine operation modes.

Claims (10)

1. An asymmetric air intake for a three-circuit engine of a supersonic aircraft, comprising a spatial wedge, a shell, side walls, a subsonic diffuser, and a throat and a boundary layer control system, where the wedge consists of compression panels, characterized in that the boundary layer control system is provided An axial separation channel is placed at the outlet of the subsonic diffuser with a wedge-shaped drainage slot located in the throat section, the front edge of which is made of complex shape, described my formula is:

where D is the diameter of the separation channel, m;

- the height of the protrusion, m; ϕ _CR - the angle of conjugation of the ledge right, deg; ϕ _lev - the angle of conjugation of the protrusion of the left, deg; ϕ _center — angle at the apex of the protrusion, degrees;

- Heaviside function. 2. The air intake according to claim 1, characterized in that the side walls are made with reverse sweep of the front edges, which are located between the compression panels of the wedge and the shell. 3. The air intake according to claim 1, characterized in that the front edge of the shell is made with reverse sweep.

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