RU2491439C1 - Turbojet protected against self-guided rocket and method of its protection (versions) - Google Patents

Abstract

FIELD: engines and pumps.SUBSTANCE: turbojet comprises housing accommodating air intake, axial-flow compressor, combustion chamber, turbine vanes and afterburner. Afterburner inner surface is coated by heat-isolation coating with reflection factor at working temperature smaller than unity. Turbine rear support cowl is shaped to direct circular or truncated cone with vertex angle exceeding 90 degrees. Jet nozzle has throat with radius smaller than that of the circle circumscribed by turbine vane ends. Turbojet may be provided with jet nozzle that has the throat with radius equal to or smaller than turbine rear support cowl base. For protection of turbojet its vanes are irradiated simultaneously by microwave-band radio waves and radio waves reflected from rear support cowl via said nozzle and rear support cowl. Note here that radio waves are irradiated outward via het nozzle with decreased amplitude after multiple reflections from heat-protection coating as well as those reflected from turbine vanes.EFFECT: decreased probability of hitting.8 cl, 4 dwg

Description

The invention relates to military equipment, and in particular to methods of individual protection of aircraft from missiles equipped with homing heads (GOS). The scope of the invention is the protection of an aircraft with a turbojet engine from missiles equipped with GOS operating in the microwave range of radio waves.

A known method of protecting aircraft from missiles equipped with homing heads (RU, patent No. 2141094, F41J 2/02, priority 17.08.1998), by creating in the space between the aircraft and the most likely direction of a possible missile attack of the enemy of false targets. This way in space as a false target, form its holographic image using a real source that emits electromagnetic waves of the visible and infrared spectra.

This method cannot provide protection for aircraft with turbojet engines in the microwave range of radio waves.

Signs of an analogue are common with the invention: receiving GOS of radio waves emitted by an object.

A known method and device for personal protection of an aircraft (LA) from guided missiles (RU 114029, IPC F41J 2/02. 03/10/12.), Which is adopted as a prototype of the method. This device contains an exhaust device mounted on the aircraft; towed radar trap mounted on the release device with the possibility of separation and connected by a towing cable to the release device. Moreover, the towed radar trap until the separation is placed on the outer surface of the release device with the possibility of switching in the mode of re-emission of guided missile (RF) signals with homing radar heads in the direction of the front and (or) rear hemispheres of the aircraft protection before separation from the release device. When a guided missile (SD) attack with a radar homing head is detected from the front or rear hemisphere, the trap is switched into the re-emission mode of the sounding signals of the SD, and then it is separated from the release device by command of the separation mechanism using locks. In the process of the trap moving away from the aircraft’s side, the UE is redirected from the signal reflected from the aircraft to the signal relayed through the receiving and transmitting antennas of the trap. After the trap leaves the aircraft for the full length of the cable and the task is completed, the trap is reset with the cable on command to the reset device.

The disadvantage of the prototype, in relation to aircraft with turbojet engines (turbojet engines) having a large effective scattering surface in the rear hemisphere, which is modeled by the sound frequency of the turbine blades, is a high probability of capture of the GSP turbojet engine, because the radar trap is only a GPS signal repeater, which does not change the spectrum of the GPS signal, which does not allow the GOS to identify reflections from the turbojet engine against the background of reflections from other objects.

Signs of the prototype are common with the invention: irradiation of aircraft with GOS radio waves and their reradiation.

Known turbojet engine (RU, patent No. 2237185, MKI F02K 3/00, priority 03/31/2003), which contains sequentially installed in the housing: an air intake, an axial compressor, a combustion chamber, a turbine with blades, a conical fairing of the back support of the turbine, afterburner combustion chamber (FCC) and jet nozzle.

When the microwave range is irradiated in the rear hemisphere of a turbojet engine (turbojet engine) by the homing head (HSP), the radio waves penetrate the FCC through the jet nozzle, are reflected from the turbine blades and the cowl of the rear turbine support, and from it in the direction of the inner side wall of the FCC, which reflects them in the direction of the jet nozzle, and the nozzle emits radio waves in the direction of the GOS, practically with the intensity of the incident waves.

Signs of an analogue that coincide with the features of the invention: an air intake, an axial compressor, a combustion chamber, a turbine with blades, a conical fairing of the back support of the turbine, an afterburner of combustion (FCC) and a jet nozzle.

Known turbojet engine (RU, patent No. 2 418 969, MKI F02K 3/02, priority 03.03.2009), adopted as a prototype of the invention, which contains sequentially installed in the housing: compressor, main combustion chamber, turbine, afterburner and reactive nozzle.

Signs of an analogue that coincide with the features of the invention: an air intake, an axial compressor, a combustion chamber, a turbine with blades, a conical fairing of the back support of the turbine, an afterburner of combustion (FCC) and a jet nozzle.

The disadvantages of the prototype device.

The prototype of the turbine blades are in the line of sight from the side of the jet nozzle, so they fully reflect the radio waves of the missile seeker falling on them.

The effective dispersion surface (EPR) of the turbine blades is hundreds of square meters, which makes it possible for the GOS to direct the missile onto the turbojet engine.

The turbine blades have an operating temperature of 1000-1300 ° C, so their EPR can not be reduced either by the use of radar absorbing material, or by changing their shape. The shape of the blades is made in strict accordance with the requirements of the laws of gas dynamics, which are different from the requirements of the laws of electrodynamics, and any radar absorbing material will burn in the hot gases of the afterburner. In addition, the turbine blades create an angular modulation of the sound frequency radio waves, which is proportional to the speed of the turbine and the number of its blades. This modulation can be used in the GOS work program to identify reflections from the turbojet engine against the background of reflections from other objects.

The fairing of the rear turbine support of the prototype has a lively - well streamlined shape (dashed lines in Fig. 2), with angles of incidence of the radio waves α of more than 45 ° from its side surface, which fall on the turbine blades, and from them at a large angle of incidence ξ are reflected in the direction of the surface of the thermal barrier coating of the FCC and after a single reflection, the PC is emitted in the direction of the GOS (figure 4).

The prototype of the invention has in the rear hemisphere a large total EPR of the turbine blades, the fairing, together with the internal side walls of the FCC, which makes it possible for the GOS to direct the missile onto the turbojet engine, which is a disadvantage of the prototype.

The objective of the invention: to reduce the likelihood of hitting aircraft with turbojet engines with missiles equipped with homing heads by reducing the effective dispersion surface (EPR) of the turbojet engine in the microwave range of radio waves in the rear hemisphere.

The technical result of the use of the invention is to reduce the likelihood of hitting aircraft with turbojet engines rockets equipped with homing heads operating in the microwave range of radio waves, by reducing the effective scattering surface of the turbojet engine in the rear hemisphere, by giving the fairing of the back support of the turbine a straight circular or truncated cone, with an angle at the apex of the cone more than 90 ° and also due to the fact that the radius of the throat of the jet nozzle is less than the radius of the circle described by the ends of the turbine blades other.

The invention is illustrated by drawings.

Figure 1 shows a diametrical longitudinal section of a turbojet engine (turbojet engine) with afterburner combustion chamber, a fairing of the rear turbine support and a jet nozzle according to the invention.

Figure 2 shows the diametrical longitudinal section (dash-dotted lines), fairings according to the invention with a cylindrical base (solid lines) and fairing prototype (dashed lines).

Figure 3 presents a diagram of the path of the rays of the radio waves incident on the fairing according to the invention with a cylindrical base, reflected from it and from the inner surface of the FCC housing, in the figure the rays are shown by lines with arrows, and dotted lines indicate the perpendiculars to the incidence planes of the GSN radio wave beam.

Figure 4 presents a diagram of the path of the rays of the radio waves incident on the fairing of the prototype, reflected from it in the direction of the turbine blades, and from them incident and reflected from the inner surface of the housing of the FCC, in the figure the rays are depicted by lines with arrows, and dotted lines indicate perpendiculars to the planes incidence of a beam of radio waves of the seeker.

Figure 1-4 introduced digital notation: 1 - air intake; 2 - axial compressor; 3 - combustion chamber; 4 - turbine; 5 - turbine blades; 6 - afterburner combustion chamber (FCC); 7 - thermal protection coating (TZP); 8 - fairing of the rear turbine support; 9 - jet nozzle (PC); 10 - subsonic shutters (DZS) of the jet nozzle, 11 - supersonic shutters (SZS) of the jet nozzle; 12 - throat of a jet nozzle (GDS).

Figure 3 and 4 introduced the legend:

α is the angle of incidence of the GOS radio waves on the line forming the side surface of the fairing of the turbine back support;

θ is the angle, at the point of incidence of the beam of radio waves, between the line forming the surface of the fairing, and the direction of the incident radio waves of the GOS missile;

β is the angle of incidence of radio waves on the thermal current transformer in the inner side surface of the FCC housing;

ξ is the angle of incidence of the radio waves on the turbine blades reflected from the side surface of the fairing of the prototype;

R is the radius of the body of the FCC;

r is the radius of the base of the fairing;

h is the height of the conical part of the fairing according to the invention;

l is the distance between adjacent points of reflection of the beam of radio waves located on a longitudinal line on one side of the inner side surface of the FCC housing according to the invention.

The technical result of the invention is that the device is achieved due to the fact that the turbojet engine (Fig. 1) is sequentially installed in the housing: air intake 1, axial compressor 2 (OK), combustion chamber 3 (KS), turbine 4, turbine blades 5, afterburner 6 (FKS), heat-shielding coating 7 (TZP), fairing 8 of the turbine back support 4, jet nozzle 9 (PC), subsonic shutters 10 (DZS), supersonic shutters 11 (SZS) and throat of the jet nozzle 12 (GDS).

The air intake 1 is intended for air intake and is made in the form of a nozzle with a cowl installed inside it of the front support of the axial compressor coaxially to the axis of the nozzle.

Axial compressor 2 (OK) is designed to create air pressure entering the combustion chamber 3 and afterburner combustion chamber 6, due to which they provide complete combustion of fuel, for example, kerosene. OK 2 contains a rotor with blades radially rigidly fixed on it, which is fixed motionless on the same axis with turbine 4.

The combustion chamber 3 (KS) is made annular of metal, and is internally coated with a heat-protective coating 7. In the chamber, air is mixed with atomized fuel, and the mixture is burned, in which high-pressure gases are formed.

The turbine 4 during operation of the combustion chamber provides rotation of the axial compressor 2. The turbine consists of a rotor with blades radially rigidly mounted on it 5. The radius of the turbine 4 rotor is equal to the radius of the radome 8. The blades 5 of the turbine are made of heat-resistant metal, the shape of which complies with the laws of gas dynamics for turbines.

According to the invention, the blades 5 of the turbine 4 perform the function of reflectors of radio waves.

Afterburner combustion chamber 6 (FCC) is designed to increase the thrust of the turbojet engine when fuel is introduced into it and burned. It is made of metal in the form of a cylindrical pipe with a skirt - a conical output device, covered from the inside with a heat-protective coating 7.

According to the invention, FCC 6 performs the function of a reflector and an absorber of radio waves.

The heat-insulating coating 7 internally covers the body of the combustion chamber (KS) 3 and afterburner combustion chamber (FCC) 6. It is made, for example, of fiberglass impregnated with soot, and at a working temperature has a power reflection coefficient less than unity, for example, 0.5 or minus 3 dB.

According to the invention TZP 7 performs the function of a radar absorbing material.

The fairing of the rear support of the turbine 8 has the shape of a round straight or truncated cone with an angle at the apex greater than 90 °, but less than 120 ° with or without a cylindrical base. The fairing is made of an alloy of metals, for example, of a heat-resistant alloy based on nickel. The fairing is installed behind the rear support of the turbine 4 at the front end of the FCC 6. The radius of the radome base is equal to the radius of the turbine rotor 4. The radius of the cylindrical radome base is equal to the radius of the base of the cone or truncated cone of the fairing 8, and its length is selected from the condition of stable operation of the turbojet engine within 3 dimensions of the height of the cone fairing. The cowl 8, made in the form of a truncated cone, the apex may be convex, for example, in the form of a spherical segment.

According to the invention, the fairing performs the function of a reflector of the incident radio waves in a direction that ensures the incidence of the radio waves at the thermal current transformer at angles greater than zero but less than 20 °.

Jet nozzle (PC) 9 is designed to impart supersonic speeds arising from the FCC, which create the thrust of the turbojet engine. PC 9 has a casing made in the form of a metal conical pipe, inside of which subsonic 10 and supersonic 11 shutters are movably mounted around the circumference, internally coated with heat-protective coating 7 (TZP). The PC case with the front end is rigidly and hermetically attached to the end of the conical output device of the FCC 6 case and subsonic shutters 10 are movably attached to it from the inside, the supersonic shutters 11 are movably attached to their free ends. The junction of the subsonic 10 and sound 11 shutters forms the throat of PC 12 ( GDS).

According to the invention, the jet nozzle 9 performs the function of a receiving and transmitting antenna, and its throat performs the function of a device completely or partially obscuring the turbine blades from GOS radio waves.

A method of protecting turbojet aircraft engines from missiles equipped with homing heads

The technical result of the use of the invention - the method is achieved in two versions of the invention.

The first embodiment of the invention is a method.

A method of protecting aircraft turbojet engines from missiles equipped with homing heads is that through the throat 12 of a jet nozzle 9 of a turbojet engine, the radius of which is less than the radius of the circle described by the ends of the blades 5 of the turbine 4, they are simultaneously irradiated with microwave waves of the microwave range of the homing head of the turbine blade rocket and the fairing 8 of the rear turbine support, which is located coaxially at the end of the afterburner and is made in the form of a cone or a truncated cone with an angle at the apex greater than 90 ° (FIGS. 2 and 3). Then, the radio waves reflected from the surface of the fairing 8 of the turbine back support are repeatedly irradiated with a heat-shielding coating, the reflection coefficient of which at the operating temperature is less than unity, for example, minus 3 dB. TZP 7 covers the inner surface of the afterburner. Then, at the same time, radio waves with a reduced amplitude, after they are repeatedly reflected from the heat-shielding coating 7, and radio waves reflected from the part of the turbine blades 5 irradiated by the radio waves, are radiated out through the jet nozzle.

These actions with the RNG radio waves of the missile, which irradiate the fairing 8 of the turbine back support, the inner lateral surface of the FCC and partially the turbine blades through the jet nozzle of the turbojet engine, reduce the total effective dispersion surface of the turbojet engine, and, consequently, also reduce the likelihood of a plane with a turbojet engine.

The second variant of the invention is a method.

A method of protecting aircraft turbojet engines from missiles equipped with homing heads is that through the throat 12 of a jet nozzle 9 of a turbojet engine, the radius of which is equal to or less than the radius of the base of the fairing 8 of the rear turbine support, which is made in the form of a cone or a truncated cone with an angle at the peak is greater than 90 ° (FIGS. 2 and 3) and is located coaxially at the end of the afterburner of the combustion chamber and is irradiated with radio waves of the microwave range of the homing missile. Then, the radio waves reflected from the surface of the fairing 8 of the turbine back support are repeatedly irradiated with a heat-shielding coating 7 covering the inner surface of the afterburner, the reflection coefficient of which at the operating temperature is less than unity, for example, minus 3 dB. Then, the radio waves with a reduced amplitude, after their multiple reflection from the heat-shielding coating 7, are radiated out through the jet nozzle 9.

These actions with missile seeker radio waves, which through the jet nozzle of the turbojet engine only irradiate the cowling of the turbine back support and the inner side surface of the FCC, significantly reduce the total effective scattering surface of the turbojet engine, and, therefore, significantly reduce the likelihood of an aircraft with a turbojet engine.

A device for implementing the method of protecting turbojet aircraft engines from missiles equipped with homing heads is made in two versions of the invention.

According to the first embodiment of the invention, the device (Fig. 1) comprises: sequentially installed in the housing: an air intake 1, an axial compressor 2, a combustion chamber 3, a turbine 4, turbine blades 5, an afterburner of combustion 6, a heat shield 7, a conical fairing 8 of the rear support turbines, a jet nozzle 9 with a throat 12. A heat-shielding coating 7 with a reflection coefficient at an operating temperature of less than unity is applied to the inner surface of the afterburner of combustion 6 and the jet nozzle 9.

The conical fairing 8 of the back support of the turbine is made in the form of a straight circular or truncated cone with an angle at the apex greater than 90 °, and the radius of the throat of the jet nozzle is less than the radius of the circle described by the ends of the turbine blades (figure 2).

The fairing 8 of the rear turbine support can be made in the form of a direct circular or truncated cone with or without a cylindrical base. The diameter of the cylindrical base is equal to the diameter of the base of the cone of the fairing, its length is less than 3 dimensions of the height of the cone (figure 3).

A truncated cone can be made with a convex peak, for example, in the form of a spherical sector.

The radius of the throat 12 of the jet nozzle 9 of the turbojet engine is smaller than the radius of the circle described by the ends of the blades 5 of the turbine 4,

According to a second embodiment of the invention, the device comprises: sequentially installed in the housing: an air intake 1, an axial compressor 2, a combustion chamber 3, a turbine 4, turbine blades 5, a combustion afterburner 6, a heat shield 7, a conical fairing 8 of the back support of the turbine, a jet nozzle 9 with a throat 12. A heat-shielding coating 7 with a reflection coefficient at an operating temperature of less than unity is applied to the inner surface of the afterburner of combustion 6 and the jet nozzle 9.

The conical fairing 8 of the back support of the turbine is made in the form of a straight circular or truncated cone with an angle at the apex greater than 90 °, and the radius of the throat of the jet nozzle is less than the radius of the circle described by the ends of the turbine blades (figure 2).

The fairing 8 of the rear turbine support can be made in the form of a straight circular or truncated cone with a cylindrical base with a diameter equal to the diameter of the base of the fairing cone, and a length of less than 3 sizes of the height of the cone (figure 3).

A truncated cone can be made with a convex peak, for example, in the form of a spherical sector.

The radius of the throat 12 of the jet nozzle 9 of the turbojet engine is equal to or less than the radius of the radome base 8.

The operation of the devices according to the invention.

During the flight of an aircraft with turbojet engines above the war zone, the aircraft may be subjected to rocket fire in the rear hemisphere, for example, ground - air from the GOS. GOS can launch a missile at a target if the sensitivity of its receiving system is higher than the signal reflected from the target.

To reduce the likelihood of an aircraft hitting a rocket in the rear hemisphere of the EPR, its turbojet engine should be as small as possible.

GOS missiles through the throat 12 of the jet nozzle 9 of a turbojet engine, at the same time irradiates with radio waves of the microwave range the radome 8 of the rear support of the turbine and partially the blades 5 of the turbine 4.

Since the turbine blades are partially irradiated by radio waves, their EPR will be less than when they are completely irradiated.

The fairing 8 reflects the SHG radio waves incident on it at a small angle of incidence on the FCC heat shield (Fig. 3), which ensures their multiple reflection from the heat shield to radiation from the jet nozzle. Therefore, the power of the radio waves irradiating the fairing, when they exit the jet nozzle, decreases by a factor of (K _TPP ) ^and times, where K is the reflection coefficient of the thermal current transformer in power at normal incidence and operating temperature; and - the number of reflections of the beam of radio waves from the FC of the FCC housing. In this case, the total EPR of the turbine and fairing blades decreases, which reduces the likelihood of a missile hit by an aircraft with a turbojet engine.

In the second embodiment of the invention, only the radome 8 and TZP 7 are irradiated by the GSP radio waves. The turbine blades are in the shadow created by the throat of the jet nozzle, therefore, they are not irradiated and do not increase the ESR of the turbojet engine, therefore this embodiment of the invention is preferred to achieve the best technical result of using the invention.

The theoretical background of the invention

Formulation of the problem

It is required to determine the angle between the intersecting straight line forming the surface of the fairing, made in the form of a cone or a truncated cone, and its axis, when the fairing is irradiated with radio waves along its axis, providing multiple reflection of radio waves from the thermal current transformer of the internal surface of the FCC housing.

EPR fairing according to the invention.

In the approximation of beam optics, formulas are obtained that calculate the required dimensions of the fairing, according to the second embodiment of the invention, when the radius of the throat PC is equal to or less than the radius of the base of the fairing.

If the radius of the throat PC is less than only the radius of the circumference described by the ends of the turbine blades, the effect of reducing the EPR of the turbojet engine in the rear hemisphere will be less.

The calculation formulas are obtained from the propagation geometry of the radiation waves of the seeker in the housing of the FCC (figure 3).

1. Determination of the optimal value of the angle of incidence β of radio waves on the heat-shielding coating (TZP) of the FCC housing, reflected from the side surface of the fairing

From the geometry of the path of the rays in the FCC (figure 3), the equalities follow:

$$2\alpha +\beta =\alpha +\theta =90°(\mathrm{one})$$

To determine the angle β, it is convenient to represent the angle θ as the sum of two angles 45 ° + φ, for φ> 0, the radio waves reflected from the side surface of the fairing do not fall on the turbine blades.

From equalities (1) follows the identity:

. $$\beta =2\phi (2)$$

.

The smaller the angle β, the greater the number of times the beam of radio waves will be reflected from the FCS FCS, providing a decrease in their intensity in the direction of PC and GOS. In order for the beam of the radio wave reflected from the TZP not to fall back on the surface of the fairing, and reflected from its surface not to go directly to the output of the PC (Fig. 3), it must be equal to or greater than the angle determined by the formula:

. $$\beta =2[45°-arctg(r/h)](3)$$

.

For the given dimensions of the fairing r and h, the optimal value of the angle β is determined by formula (3), which ensures the minimum intensity of the beam of radio waves that returns to the PC, and from it in the direction of the GPS, due to the fact that the beam reflected from the TZ is not returned on a fairing, but it is repeatedly reflected from the TZP.

2. Determination of the distance 1 between adjacent points of reflection of the beam of radio waves located on a longitudinal line on one side of the inner side surface of the FCC housing.

We write the formula following from the construction of the rays of the radio waves in figure 3:

. $$\mathrm{one}=\mathrm{four}R\cdot tg\beta (\mathrm{four})$$

.

3. Determination of the EPR (σ _min ) FCC with a fairing of the optimal electrodynamic shape according to the invention.

For physical reasons, the formula for determining σ _{min is} written as:

$${\sigma }_{\min }=\sigma \cdot {(\mathrm{TO}tPP)}^{\mathrm{but}}(5)$$

σ _min - EPR FKS with fairing optimal electrodynamic shape in;

σ - ESR FKS with fairing prototype;

K is the reflection coefficient of the thermal current transformer in terms of power at normal drop and operating temperature;

and - the number of reflections of the beam of radio waves from the TZP of the FC housing, which is determined by the formula (6):

$$a=2\cdot L/\mathrm{one}(6)$$

L is the length of the body of the FCC;

l is the distance between adjacent points of reflection of the beam of radio waves from the TZ located on a longitudinal line on one side.

Calculation of ESR reduction in turbojet engine in the back hemisphere

1. EPR Evaluation FCC fairing prototype. The fairing is made of metal in the form of a round cone with a surface of a lively shape that satisfies the laws of gas dynamics, but does not take into account the requirements of electrodynamics, therefore, it cannot minimize the ESR of a turbojet engine, because Radio waves fall and are reflected from the fairing surface at large angles at which the reflection coefficient of any materials tends to unity. The reflected radio waves from it fall onto the turbine blades at small angles of incidence and, with no more than a single reflection from the inner surface of the PCS at a large angle of incidence - reflection (Fig. 4), fall into the throat of the jet nozzle and are emitted in the direction of the seeker.

The ESR of the fairing of the prototype a installed in the FCC, when the radio waves are reflected from its lateral surface towards the turbine blades, and from them in the direction of the PC and GOS, is determined by the formula (7) of the ESR for a circular disk in the direction of its axis:

$$\sigma =\pi \cdot {(2\pi /\lambda )}^2\cdot {(r^2)}^2(7)$$

,

where λ is the working wavelength;

r is the radius of the radome base.

EPR fairing of the prototype, with a radius of the base, for example, r = 0.22 m at a wavelength of λ = 0.02 m is 180 m ² .

2. Evaluation of the EPR FCC fairing according to the invention.

EPR FCC with a fairing according to the invention - σ _min calculated by the formulas (2), (4), (5) and (6).

The initial data for the calculation:

R = 1.0 m, r = 0.4 m, L = 1.4 m, K = 0.5, σ = 180 m ² .

Determination of the angle θ. Set the angle ϕ, which should be greater than zero, for example, 3 ° and determine the value of the angle θ by the formula θ = 45 ° + ϕ = 46 °,

By the formula (2) determine β = 2ϕ = 6,

By the formula (4), the value l = 4R · tgβ = 4 · tg6 ° = 0.42 m is determined.

By the formula (6) determine the value $$a=2\cdot L/\mathrm{one}=\mathrm{2,8}/0.42=6.66~7(6).$$

By the formula (5) determine the value of σ _min :

, $${\sigma }_{\min }=\sigma \cdot {({\mathrm{TO}}_{tPP})}^{\mathrm{but}}=180{\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="00000011.png"\; state="\{\{state\}\}">m</figure-callout>}}^2\cdot {(0.5)}^{\mathrm{but}}=180{\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="00000011.png"\; state="\{\{state\}\}">m</figure-callout>}}^2\cdot 0.0078=1.4m^2(5)$$

,

The decrease in the EPR of the FCC with the fairing according to the invention in dB, compared with the fairing of the prototype is determined by the formula:

101g (σ / σ _min ) = 101g (180 m ² / 1.4 m ² ) = 101g128.6 = 10 · 2.11 = 21 dB.

Conclusion

The EPR FKS with a fairing of the optimal electrodynamic shape is less than the EPR FKS with a fairing of the prototype by 21 dB.

The technical result of the invention is achieved, because with a decrease in the EPR of the turbojet engine in the rear hemisphere by 21 dB, the likelihood of hitting aircraft with turbojet engines with missiles equipped with homing heads operating in the microwave range of the radio waves decreases. Distinctive features of the claims.

The first embodiment of the invention is a method.

Before irradiation, a heat-shielding coating with a reflection coefficient at an operating temperature of less than unity is applied to the inner surface of the afterburner chamber of the turbojet engine.

The radius of the throat of a jet nozzle of a turbojet engine is less than the radius of the circle described by the ends of the turbine blades, which are mounted coaxially with the fairing of its rear support.

The fairing is made in the form of a cone or a truncated cone with an angle at the apex greater than 90 °.

The radio waves of the homing head carry out simultaneous irradiation, through the aforementioned throat of the jet nozzle, turbine blades and the cowling of its rear support.

Then, the radio waves reflected from the surface of the fairing of the rear turbine support are repeatedly irradiated with said heat-shielding coating, while the radio waves with reduced amplitude, after being repeatedly reflected from the heat-shielding coating, and the radio waves reflected from the said turbine blades, are simultaneously radiated outward through the jet nozzle.

The second variant of the invention is a method.

Before irradiation, a heat-shielding coating with a reflection coefficient at an operating temperature of less than unity is applied to the inner surface of the afterburner chamber of the turbojet engine.

The radius of the throat of a jet nozzle of a turbojet engine is equal to or less than the radius of the radome base of the rear turbine support.

The fairing is made in the form of a cone or a truncated cone with an angle at the apex greater than 90 °.

Then, the radio waves of the homing head carry out simultaneous irradiation, through the aforementioned throat of the jet nozzle, turbine blades and the cowling of its rear support.

Then, the radio waves reflected from the surface of the fairing of the rear turbine support are repeatedly irradiated with said heat-shielding coating, while the radio waves with reduced amplitude, after being repeatedly reflected from the heat-shielding coating, and the radio waves reflected from said turbine blades, are simultaneously radiated outward through the jet nozzle.

The first embodiment of the invention is a device.

The thermal barrier coating has a reflection coefficient at an operating temperature of less than one. The conical fairing of the rear turbine support is made in the form of a straight circular or truncated cone with an angle at the apex greater than 90 °. The radius of the throat of the jet nozzle is less than the radius of the circle described by the ends of the turbine blades. In addition, the conical fairing of the rear turbine support is provided with a cylindrical base, the diameter of which is equal to the diameter of the base of the fairing cone, and the length is less than three dimensions of the cone height.

The truncated cone of the fairing of the rear turbine support is made with a convex apex.

The second embodiment of the invention is a device.

The thermal barrier coating has a reflection coefficient at an operating temperature of less than one. The conical fairing of the rear turbine support is made in the form of a straight circular or truncated cone with an angle at the apex greater than 90 °. The radius of the throat of the jet nozzle is equal to or less than the radius of the radome base of the rear turbine support.

In addition, the conical cowling of the rear turbine support is provided with a cylindrical base, the diameter of which is equal to the diameter of the cone base of the cowling of the rear turbine support, and the length is less than three cone height sizes.

The truncated cone of the fairing of the rear turbine support is made with a convex apex.

Claims (8)

1. Aircraft turbojet engine, made with the possibility of protection against a missile equipped with a homing head operating in the microwave range of radio waves, comprising a housing in which an air intake, an axial compressor, a combustion chamber, a turbine, turbine blades, and a combustion afterburner are installed, the inner surface of which covered with a heat-shielding coating with a reflection coefficient at an operating temperature of less than unity, a cowl around the back support of the turbine, made in the form of a direct circular or truncated whisker with an apex angle greater than 90 ° C, and a jet nozzle with a throat having a radius smaller than the radius of the circle described by the ends of the turbine blades. 2. The turbojet engine of the aircraft according to claim 1, characterized in that the fairing is made in the form of a straight circular or truncated cone with an angle at the apex greater than 90 ° and is equipped with a cylindrical base, the diameter of which is equal to the diameter of the base of the cone of the cowling of the back support of the turbine, and the length is less than three cone height dimensions. 3. The turbojet engine of the aircraft according to claim 1, characterized in that the truncated cone of the fairing of the rear turbine support is made with a convex top. 4. A turbojet engine of the aircraft, made with the possibility of protection against a missile equipped with a homing head operating in the microwave range of radio waves, comprising a housing in which an air intake, an axial compressor, a combustion chamber, a turbine, turbine blades, and a combustion afterburner are installed, the inner surface of which covered with a heat-shielding coating with a reflection coefficient at an operating temperature of less than unity, a cowl around the back support of the turbine, made in the form of a direct circular or truncated whisker with an apex angle greater than 90 ° C, and a jet nozzle with a throat whose radius is equal to or less than the radius of the turbine back shroud base support. 5. The turbojet engine of the aircraft according to claim 4, characterized in that the fairing of the back support of the turbine is made in the form of a direct circular or truncated cone with an angle at the apex greater than 90 ° and is equipped with a cylindrical base, the diameter of which is equal to the diameter of the base of the cone of the cowling of the back support of the turbine, and less than three dimensions of the height of the cone. 6. The turbojet engine of the aircraft according to claim 4, characterized in that the truncated cone of the fairing of the rear turbine support is made with a convex apex. 7. A method of protection against a rocket equipped with a homing head operating in the microwave range of radio waves of an airplane turbojet engine, characterized in that the turbojet engine is made with a throat radius of the jet nozzle smaller than the radius of the circle described by the ends of the turbine blades and with a fairing of the rear turbine support located coaxially at the end of the afterburner of the combustion chamber in the form of a cone or a truncated cone with an angle at the apex greater than 90 °, and heat protection is applied to the inner surface of the afterburner of the engine combustion a coating with a reflection coefficient at a working temperature of less than one, while providing simultaneous exposure to microwave waves through the aforementioned throat of the jet nozzle of the turbine blades and the cowling of the rear turbine support, and by radio waves reflected from the surface of the cowling of the rear turbine support - multiple irradiation of said heat-shielding coating, and through the jet nozzle provides simultaneous emission of radio waves with a reduced amplitude after their repeated reflection from the heat-shielding and covering of radio waves reflected from said turbine blades. 8. A method of protection against a rocket equipped with a homing head operating in the microwave range of radio waves of an airplane turbojet engine, characterized in that the turbojet engine is made with a throat radius of the jet nozzle equal to or less than the radius of the radome of the radome of the rear turbine support and with the radome of the turbine back support located coaxially at the end of the afterburner, in the form of a cone or a truncated cone with an angle at the apex greater than 90 °, and heat is applied to the inner surface of the afterburner of the engine a protective coating with a reflection coefficient at a working temperature of less than unity, while providing exposure to microwave waves through the aforementioned throat of the jet nozzle of the cowling of the rear turbine support, and by radio waves reflected from the surface of the cowling of the rear turbine support - multiple irradiation of the said heat-protective coating, and through the jet nozzle provide radiation outward of radio waves with a reduced amplitude after their repeated reflection from the heat-shielding coating.

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