RU2271305C1 - Light supersonic multi-purpose aircraft - Google Patents

Abstract

FIELD: aeronautical engineering.

SUBSTANCE: proposed aircraft is provided with main power plant including supersonic engines, additional power plant located in lower section of fuselage behind crew cabin and integral complex of onboard radio electronic equipment , aircraft complex control system, shipboard equipment control system, aircraft complex power supply system, electronic display complex and self-contained onboard oxygen system. Proposed aircraft may be used under varying flight conditions including conditions of using autonomous home airfield in case of separation from main airfield.

EFFECT: enhanced efficiency.

3 cl, 15 dwg

Description

The invention relates to aircraft, in particular to the layout and equipment of the aircraft, increasing its functionality, including in isolation from the main airfield, expanding its capabilities for use at subsonic and supersonic flight speeds, and can be used to create multi-purpose aircraft, who have increased demands on maneuverability and flight speed, including when creating training aircraft.

Known maneuverable training aircraft Yak-130 (see RF patent No. 2144885 from 01/27/2000).

The main disadvantage of the known aircraft is the relatively low transport efficiency, insufficient flight performance, lack of structural elements necessary to perform the functions of a multi-purpose aircraft, namely:

lack of universal devices on the wing for fastening external suspended objects;

the absence of an auxiliary power plant that allows for autonomous launch under basing conditions in isolation from the main airfield;

lack of shutter valves in the air intake.

The closest technical solution both in design and in tasks performed, selected as a prototype, is a maneuverable light multi-purpose aircraft (see RF patent No. 2210522 C1 of 08.20.2003).

The main disadvantage of the known aircraft are insufficient flight performance, the lack of structural elements necessary to perform the functions of a supersonic multipurpose aircraft, namely:

the absence of the nose for flow around a supersonic stream;

lack of elongated influx of wing;

lack of reverse sweep of the trailing edge of the wing;

lack of slotted discharge of the boundary layer of the air intake;

the lack of an extended rear fuselage to accommodate supersonic engines of the main power plant;

the inability to place an auxiliary power plant along the axis of the aircraft in the central part of the fuselage;

lack of differential-controlled stabilizers of horizontal plumage;

lack of a universal device for mounting external suspended objects on an aircraft;

lack of an integrated complex of avionics;

lack of an integrated aircraft control system;

lack of a control system for general aircraft equipment;

lack of electronic indication complex;

lack of an integrated aircraft power supply system;

lack of an autonomous on-board oxygen system.

The technical results obtained from the use of this invention include ensuring the efficient use of a light supersonic multipurpose aircraft for flying under various conditions, including autonomous basing in isolation from the main airfield, by eliminating the above-mentioned disadvantages of the known technical solution.

The problem is solved, and the technical result is achieved by the fact that in an aircraft containing a fuselage with a cockpit in the bow, a brake flap in its upper part, side air intakes bounded by a wing, a landing gear, a main power plant with two turbojet engines located in the middle of the fuselage , swept wing with a front root influx of a lively shape, mechanization and deflectable socks with ledges of the leading edge, one-keel vertical tail and horizontal tail, according to Bretenoux, the root of the wing is formed with an influx scope l _n is not less than 0.15 and not more than 0.18 l _k console span of the wing from the base of the influx and the chord b _n is not less than 0.55 and not more than 0.6 magnitude local chord b _k wing, ledge of the leading edge of the deflected nose of the wing is made at a distance of 0.55-0.65 of the magnitude of the span l _{k of} the wing console from the base of the influx with the chord b _n not less than 0.075 of the local wing, without taking into account the ledge,

the wing has a reverse sweep of the trailing edge of the wing and is equipped with universal devices for mounting external suspended objects placed symmetrically on each of its console,

horizontal tail is made in the form of differentially controlled stabilizers, the rotation axis of which intersects with the axis of the engines of the main power plant, and the trailing edge protrudes beyond the end of the nozzle apparatuses,

contains supersonic engines of the main power plant, located in the rear of the fuselage so that the root of the horizontal tail is between them, for which the tail of the fuselage is made expanded;

contains an auxiliary power unit located at the bottom of the middle part of the fuselage behind the cockpit so that its longitudinal axis is directed along the axis of the aircraft, and the exhaust pipe is led out under the fuselage to the left side,

contains air intakes made in such a way that the inlet is located in the lower part of the fuselage, the air intake channel in the front part is inclined upward, and then smoothly goes into the horizontal position at the engine level; the upper edge of the air intake and the wing has a slit drain of the boundary layer,

contains an integrated complex of avionics (IR avionics), which includes:

information exchange system,

on-board digital flight control and combat training control system,

external storage device and information input system,

inertial system

short-range navigation and landing radio engineering system and air traffic control and state recognition transponder,

automatic radio compass

radio altimeter with transceiver and antenna device,

marker receiver

an integrated aircraft control system with system controls installed in the cockpit and operator’s cabin,

armament control system with remote controls installed in the cockpit of the pilot and operator,

integrated system of electronic indication, control and aiming,

alarm bulletin boards installed in crew cabins,

satellite communications system

double redundant control system for general aircraft equipment,

airborne objective control system,

connected radio station

airborne intercom module,

power supply system

external and internal lighting equipment,

integrated emergency exit system,

double redundant electronic powertrain control system,

contains an integrated control system (KSU) of the aircraft, which includes:

four-time redundant digital computing part with power and control units, sensors and control panels,

electro-hydraulic steering drives four times redundant in the electrical part and twice in the hydraulic part, two of which are installed in the roll channel and one in the direction channel,

four times redundant in the electrical part and twice in the hydraulic part of the electro-hydraulic steering gear of the stabilizer,

rotary electro-hydraulic flap drive,

three-position hydraulic wing sock drive cylinders,

electro-hydraulic control valves for deflectable wing socks,

brake flap drive,

control posts of the first and second cab, mechanically interconnected,

mechanisms of trimmer effect on pitch, roll and direction channels,

KSU control panels,

flexible flap control shafts,

planetary gears-hinges for flaps

brake flap control system,

contains a control system for general aircraft equipment (SU CCA), which includes:

control panel (PU CCA),

light-signal board with buttons-lamps of the PKN type (TS PKN),

Central light-signal light (ЦСО),

multiplex information exchange channel,

block of transformations and calculations (BPV),

control and monitoring unit (BUK),

airborne equipment systems (CCA),

air conditioning system (SCR),

integrated management system (KSU),

multifunctional indicators (MFCI),

inter-aircraft intercom (MSPU),

pneumohydrosystem (ASG),

wheel braking system (STK),

fire protection system (SDR),

lantern visor glass air protection system (SVZS),

Emergency Escape Facility (KSAP),

power plant start-up, control and monitoring system (ESA),

power supply system (SES),

fuel management and control system (SUIT),

start-up system and control of the auxiliary power unit (APU),

oxygen system (KS),

chassis control system (SS),

radio altimeter

radio compass (ARC),

lantern lock alarm system (SSZF),

external lighting equipment system (ESPO),

fuel system (TS),

blinds shutter control system (USVZ),

centralized oil filling system (SPM),

guidance and landing equipment,

rangefinder,

on-board digital computer (BTsVM)

contains an integrated aircraft power supply system, which includes:

AC subsystem consisting of:

left drive generator (GPZ 1),

right drive generator (GPZ 1),

APU generator (P / N 20032),

airdrome power source connection unit (SHRAP-400-3F),

AC bus (115 V, 400 Hz) left and right side,

contactors (K1), (K3), (K11), (K10), (K2),

regulation, protection and control units (BRZU 15 VO 3 ser.),

phase sequence unit (BChF-208),

blocks of current sensors (BDT16K and DPCT P / N 20732),

left rectifier device (VU) (VU6BKF),

right WU (VU6BKF),

DC subsystem consisting of:

left battery (KSX25P-A-TCC),

right battery (KSX25P-A-TCC),

emergency bus DC (27 V) of the left and right sides,

switchable DC bus (27 V) left and right side,

contactors (K2), (K6), (K5), (K1), (K4), (K3), (K2),

VU protection and control devices (AZUVU200B),

APU electric starter,

contains a complex of electronic indications, which include:

first managing computer,

multiplexed information exchange channel (MKIO),

local circular channel of information exchange (LKIO),

left, central and right MFCI,

kallimotor aviation indicator (KAI),

multifunctional control panels (MFPU),

power supply system with a generator of the main power plant (SU), a generator of the auxiliary power plant (APU),

accumulator battery,

sensor system of operational parameters,

second managing computer,

inter-machine channels of information exchange,

the first (left) and second (right) local radial channels of information exchange are redundant,

television signal switching unit (BKTS),

contains a universal device for mounting external suspended objects on an airplane, which includes:

pylon with wing mounts,

lockless suspension units

holder lock

compartments for placement of blocks for disconnecting and / or launching objects,

hatch

removable tail fairing,

contains an autonomous on-board oxygen system, which includes:

anti-loading suit

ejection system oxygen system,

breathing air sensor

check valve

oxygen hose

pressure machine

pressure regulator

oxygen hoses

oxygen device

oxygen mask,

protective helmet,

cab valve

equalization hose

airborne oxygen production plant

hoses

air preparation device.

The invention is illustrated by drawings.

Figure 1, 2, 3 shows schematically a General view of the aircraft, respectively, a side view, a plan view, a front view;

figure 4 - layout of the aircraft and the location of the main and auxiliary power plant;

figure 5, 6 is a layout of the main power plant, respectively, rear view and top view;

Fig.7 is a design of the drain gap of the boundary layer of the air intake and the placement of the auxiliary power unit;

on Fig is a diagram of an integrated complex of avionics;

Fig.9 is a diagram of an integrated aircraft control system;

figure 10 is a diagram of a control system for aircraft equipment;

11 is a diagram of an alternating current subsystem of an integrated power supply system;

in Fig.12 is a diagram of a DC subsystem of an integrated power supply system;

Fig.13 is a diagram of a complex of electronic indication;

on Fig - scheme of the suspension of the payload;

on Fig is a diagram of an autonomous on-board oxygen system.

The aircraft contains the fuselage 1 with the cockpit 2 in the bow, the brake flap 3 in its upper part, the side air intakes 5, the landing gear bounded by the wing 4, the landing gear 6, the main power plant with two turbojet engines located in the middle of the fuselage 7, the swept wing 4 has front root influxes of an animated shape 8, mechanization 9 and deflectable socks 10 with ledges of the leading edge 11, single-tail vertical tail 12 and horizontal tail 13, root influx 8 of wing 4 is made with a span of l _{n of} not less than 0.15 and not more than 0.18 xa l _k the wing console 4 from the base of the influx 8 and with the chord b _n not less than 0.55 and not more than 0.6 of the local wing chord b _k , ledge 11 of the leading edge of the deflected nose 10 of the wing 4 is made at a distance of 0.55-0 , 65 the magnitude of the span of the wing console l _k from the base of the influx 8 with the chord b _{n of} at least 0.075 the magnitude of the local chord b _{k of} wing 4 without taking into account step 11, wing 4 has a reverse sweep of the trailing edge of the wing and is equipped with universal devices 14 for mounting external suspended aircraft 15 objects placed symmetrically on each of its console, horizontal fencing 13 is made in the form of differentially controlled stabilizers, the axis of rotation of which intersects with the axis of the engines 7 of the main power plant, and the trailing edge protrudes beyond the end of the nozzle apparatus 16, supersonic engines 7 of the main power plant are located in the rear of the fuselage 1 so that the root part of the horizontal the plumage 13 is located between them, for which the rear part of the fuselage 1 is made extended, auxiliary power unit 17 (APU), located at the bottom of the middle part of the fuselage 1 behind the cockpit 2 so that its longitudinal axis is directed along the axis of the aircraft, and the exhaust pipe 18 is led out under the fuselage 1 to the left side, the air intakes 5 are made so that the inlet 19 is located at the bottom of the fuselage 1, the channel 20 of the air intake 5 in the front part is inclined upward, and then at engine level 7 smoothly goes into a horizontal position, in the upper part of the air intake 5 limited by the wing 4, controlled slit shutters 21 are provided with air recharge, a drain slot 22 is made between the upper edge of the air intake 5 and the wing 4 boundary layer.

The integrated complex of on-board electronic equipment 23 (IC BREO) contains an on-board automated control system 24, an on-board digital computer system 25 (BTsVS) with on-board digital computers 26, 27, an external storage device (VZU) 28, an integrated electronic indication system 29 (KSEI) , sighting and flight indicator (PPI) 30, multifunctional digital indicators (MFTS) 31-33, multifunctional control panels (MFPU) 34, switching block of television signals (BKTS) 35, voice information -controlling system (RIUS) 36, information input system (SVI) 37, helmet-mounted target designation system (NSC) 38, system multiplex information exchange channels (SMKIO) 39-41, strapdown inertial navigation system (SINS) 42, integrated with satellite navigation system 43, short-range navigation and landing radio engineering system (RSBN) 44, automatic radio compass (ARC) 45, marker radio receiver (MCI) 46, air traffic control (ATC) 47, radio altimeter (RVM) 48, airborne electronic complex (BREC) 49, onboard front-view radar station (BRL) 50, rear-view radar station (BRL) 51, on-board digital computer system (BCVS) 52, optoelectronic aiming system (OEPK) 53, optoelectronic aiming system (OEPS) front hemisphere 54 ( in a hanging container), survey-tracking heat direction finder (OSTP) of the rear hemisphere 55, small-sized thermal imaging system (TPS) 56 (in a hanging container), laser spot detector 57, recognition interrogator (ZGO) 58, local multiplex information channel exchange (LMKIO) 59, airborne defense complex (BKO) 60, electronic countermeasures complex (KREP) 61, airborne digital computer (BTsVM) 62, receiving radio intelligence modules with antennas (PRMRR) 63, transmitting radio interference modules with antennas (PMR) 64, rear-view equipment (AZO) 65, radiation warning station (SPLO) 66, technical equipment for setting curtains 67, integrated airborne communications system (IBKS) 68, radio communication module in MV-DMV range 69, communication TLK module in MV- UHF band 70, satellite module communication 71, equipment ZAS-TLF 72, equipment ZAS-TLK 73, equipment for internal communication and switching (AVSK) 74, specialized digital computer (SCV) 75, local multiplex information exchange channel (LMKIO) 76, integrated aircraft control system (KSU) 77 with system consoles installed in the crew’s cabin, weapons control system (FCS) 78, on-board digital computer (BTsVM) 79, local multiplex information exchange channel 80, interface units with LMKIO (OU SUO) 81, executive units (BI) 82, providing unloading external external objects 15, including emergency, devices for the ejection of powder cartridges (HC) 83, a control system for general aircraft equipment (SU CCA) 84, an objective control system (SOC) 85, an on-board recording device (BUR) 86 with a collection and processing unit digital and parametric information (BSPI), protected on-board drive (ZBN), operational on-board drive (EBN), television objective control system (T-SOK) 87, cameras in the cabin space and video recorder, automatic control system and power plant control system (ACS SU) 88, 89, low-speed engine control handle (ORE) in crew cabins 2, aircraft emergency exit system (KSAPS) 90, power supply system (SES) 91, generator 92 on the APU 17, onboard maintenance panels 93.

The integrated control system (KSU) 77 contains a four-time redundant digital computing part 94 with power and control units, sensors and control panels, four-time redundant for the electrical part and two-time for the hydraulic part electro-hydraulic steering gears 95 (RPD-15A), two in the roll channel and one in the directional channel, four times redundant in the electrical part and two times in the hydraulic part, the electro-hydraulic steering stabilizer drive 96 (RPD-17A), rotational electro-hydraulic wing flaps 97 (RPZ-130), three-position hydraulic actuator cylinders 98 (three-position actuator cylinders 98 deflectable socks 10 of wing 4 are located in wing 4 in the area of the first spar perpendicular to the axis of rotation of socks 10 and are directly connected to them) of wing socks 10, electro-hydraulic taps control (EGC) 99 deflectable socks 10 of the wing 4, the drive 100 of the brake flap 3, the control posts 101 of the first and second cab, mechanically interconnected, the mechanisms of the trim effect 102 through the channels of pitch, roll and direction (MP-10S6A14), control panels 103 of KSU (from the control panel of KSU there is a manual switch to backup software in case of failure of the main one), flexible shafts (not shown in Fig. 6) for controlling flaps 104, planetary gearboxes-hinges 105 flaps 104 (RSh-43), the control system 106 of the brake flap 3, the actuators 107 ailerons 108 located along the third side member of the wing 4 and connected with the ailerons 108 through the rocker 109 and rod 110, the hinge gear 111, driven from the drive located in the fuselage at help flexible shafts installing enes on the fifth wing spar 4 and flaps 104 are connected with the driving levers 112 and rods 113.

The universal control system for general aircraft equipment (OS OCO) 84 (Fig. 7) contains a control panel (PU OCO) 114, a signal board with buttons-lamps of the PKN type (TS PKN) 115, a central light-signal light (CSO) 116, a multiplex communication channel (MKIO) 39, block of transformations and calculations (BPV) 117, control and monitoring unit (BUK) 118, general aviation equipment systems (CCA) 119, air conditioning system (SCR) 120, integrated control system (KSU) 77, multi-functional indicators ( MFCI) 31, inter-aircraft transportation on-air device (MSPU) 121, pneumatic and hydraulic system (ASG) 122, wheel braking system (STK) 123, fire protection system (SPZ) 124, air protection system for the lamp visor glass (SVZS) 125, emergency exit system (KSAP) 90, system start-up, control and monitoring of the power plant (ESA) 126, power supply system (SES) 91, fuel control and control system (SUIT) 126, start-up and control system of the auxiliary power plant (APU) 127, oxygen system (CS) 128, control system chassis (school) 129, radio altimeter 48, radio compass (ARC) 45, system topic of lantern lock signaling (SSZF) 130, external lighting equipment system (ESPO) 131, fuel system (TS) 132, blinds shutter control system (USVZ) 133, centralized oil filling system (SPM) 134, guidance and landing equipment 135, range finder 136 , BTsVM 26, 27.

SU OSO 84 is interfaced via multiplex channel 39 with a complex of on-board digital computers 26, 27, an electronic control system for the left and right engine 126, a registration and control system, an integrated control system 77. And along code lines of communication (CLS) with a control and control system fuel 126, voice communications equipment 121, an integrated electronic display system 31, and auxiliary power unit 127. SU OCO 84 has two control loops structurally designed into primary and backup BPV 117 units operating in "hot standby" connected to the executive devices through BUK 118.

The integrated power supply system 91 of the aircraft (FIGS. 8, 9) contains, in the AC subsystem (FIG. 8), a left drive generator 137 (GPZ 1), a right drive generator 138 (GPZ 1), an APU generator 139, an airfield source connection unit electricity 140 (SHRP), AC bus (115 V, 400 Hz) left 141 and right 142 sides, contactors K1 143, K3 144, K11 145, K10 146, K2 147, control, protection, control units 148-150 (BRZU ), phase rotation unit 151 (BChF), current sensor blocks 152-154 (BDT), left rectifier device 155 (VU) and right 156 (VU), in the DC subsystem left VU 155, right VU 156, left rechargeable battery 157, right rechargeable battery 158, emergency DC bus (27 V) left 159 and right 160 boards, disconnected DC bus (27 V) left 161 and right 162 boards, K2 contactors 163, K6 164, K5 165, K1 166, K4 167, K3 168, K2 163, protection and control devices VU 169 and 170, electric starter APU 171.

The integrated electronic display system 29 contains (FIG. 10) a first control BCMC 26, a multiplex information exchange channel (MKIO) 39, a local circular channel of information exchange (LCIO) 172, the left, central and right MFCI 31, 32, 33, KAI 173, MFPU 34, power supply system 91 with the generator of the main power plant 137, with the generator of the APU 139, with the battery 157, the sensor system of operational parameters 174, the second control computer 27, inter-machine communication channels 175, the first (left) and second (right) local radial ca ala redundant information exchange 176, 177, the switching unit television signals (BKTS) 35.

A universal device for mounting external hanging objects on an aircraft contains (Fig. 11) a pylon 14 with wing mounts 178, lockless suspension assemblies 179, holder lock 180, compartments for accommodating objects 18 and / or launch detachment blocks, hatch 182 and a removable tail fairing 183.

The onboard autonomous oxygen system 128 (FIG. 12) includes an anti-loading suit 184, an ejection system oxygen system 185, a breathing air sensor 186, a check valve 187, an oxygen hose 188, a pressure switch 189, a pressure regulator 190, an oxygen hose 191, 192, an oxygen device 193, an oxygen mask 194, a protective helmet 195, a cabin valve 196, an equalization line hose 197, an onboard oxygen production unit (BKDU) 199, hoses 200, an air preparation device 201.

A light supersonic multipurpose aircraft (LSMS) is a normal aerodynamic aircraft with a mid-swept wing, a single-tail vertical tail and an all-turning stabilizer consisting of two independent surfaces. The wing mechanization includes controlled socks (two sections on each console) and single-slotted flaps with a low degree of extension.

The layout of the aircraft implements flight performance characteristics that allow you to master maneuvers similar to those performed on aircraft of 4-5th generation.

The layout with developed influxes of the wing provides the possibility of a stable and controlled flight at angles of attack of at least 25 ° with a hydromechanical control system and a stability improvement system (CMS) and at least 30 ° with a remote control system (CDS). SLMS has a hydromechanical system in pitch, heading and roll channels with a stability improvement system.

The cockpit is single or double. Ejection seats provide crew rescue over the entire range of speeds and altitudes, including at V = 0 and H = 0. In a double cabin, the crew is located one after another. The possibility of piloting the aircraft by one pilot from the front cockpit is provided. The information and control field of the cockpit, instruments, control panels are identical to aircraft of 4-5 generations. The equipment allows you to simulate the operating mode of airborne sighting and survey systems, the modes of preparation and use of the main types of guided weapons of 4-5th generation aircraft.

A brake flap is installed behind the aircraft cockpit, which ensures effective braking in maneuvering modes over the entire speed range, as well as possible control of pre-landing planning.

The aircraft’s hydraulic system consists of two autonomous hydraulic systems: a booster and a general one, each of which has its own pressure sources of equal power. In the booster hydraulic system, an emergency pumping station with a direct current electric motor is installed as an emergency source of pressure.

The fuel system provides power to the engines in all flight modes, including flight at near-zero overloads for at least 5 seconds, and for negative ones - 20 seconds at maximum after-load engine operation modes. The system provides centralized fueling.

The aircraft has a fire extinguishing system consisting of fire extinguishers and a sensor system.

The air conditioning system provides the required excess pressure in the cockpit in all flight modes. In the Earth’s small gas mode, the pressure and temperature in the cabin are provided by an auxiliary power unit (APU). The aircraft is equipped with an anti-icing system for the visor of the cockpit lantern.

The power plant of the aircraft includes two afterburner engines and an auxiliary power unit (APU). Right and left engines are interchangeable with each other. The power plant has an autonomous launch system from the APU. Unregulated air intakes are designed to provide sufficient performance at M <1.0, including large angles of attack, as well as minimizing traction loss at a flight speed of M = 1.4.

The design allows to provide an aircraft resource of 10,000 flight hours and 30 years of operation.

Normal LSMS aerodynamic configuration, approaching integral with the wing moderate elongation λ = 3,67, narrowing η = 3,4, swept along the leading edge χ _s.c. = 31 °, along the trailing edge χ _w.s. = -5 °, with a step along the leading edge, with a fully rotatable horizontal plumage, a single-keel vertical plumage and two afterburner engines located in the fuselage.

The aerodynamic similarity to modern airplanes is ensured by the use of “useful flow separation” realized by root wing inflows of a Gothic shape, proximity to the wing shape in plan, deflected wing socks and the use of relatively low thicknesses of bearing surfaces to ensure achievement of moderate supersonic numbers M.

To ensure the operability of the control system, a linear dependence of the longitudinal moment on the angle of attack m _z = f (α) is realized.

The relative mass of the LSMS glider is close to m _{con =} 0.31.

The combination of high maneuverability, stable and controlled flight at large operational angles of attack (α ° = 25-30) and supersonic flight speeds make the LSMS comparable and / or superior to promising T-50 and Mako aircraft.

The mass excellence of LSMS is (7-12)% higher than the Hawk-100 and Alpha Jet and commensurate with the T-4, in the construction of which composite materials are used.

The aircraft in a two-seat version with cockpits located one after the other, has a rear cockpit offset in height to provide the necessary forward view. Cabins have a common lantern opening sideways. Both cabins are equipped with ejection seats, allowing you to leave the plane at any height with an indicator speed of 0 to 1200 km / h.

Two engines are placed side by side in the rear of the fuselage, engine air intakes are located on the sides of the fuselage under influxes and have separate "S" -shaped channels to enter the engines. The channels are equipped with combined controlled shutters for additional make-up and pressure bypass at supersonic speeds such as blinds. The design of the air intake provides drainage of the boundary layer that occurs on the side surface of the fuselage and the lower surface of the influx.

Tricycle landing gear, with front controllable strut, retractable into the fuselage. The main bearings are equipped with wheels with high energy intensity of the brakes and built-in fans for cooling the brakes.

The fuselage of the semi-monocoque type with an integral joint with the upper surface of the wing has two longitudinal beams on both sides, emanating from the wing for mounting stabilizers that provide the necessary aerodynamic shoulders. On-board electronic equipment is located in the nose of the fuselage, in the cockpits, on the sides of the cabins and in the cockpit of the fuselage. Behind the second cabin, below, SCR units and an auxiliary power unit are installed that provide autonomous starting of the main engines.

In the middle part there is a fuel compartment, engine channels and a compartment for the main chassis supports, as well as a compartment for SKV and AFU units. The wing consoles and the brake flap are attached to the middle part.

The sections of the middle part of the fuselage have upper arcuate contours smoothly conjugated with the influx of the wing and directly with the upper surface of the wing; under the influx, the oval contours of the air intake pass into the contours of the fuselage.

The cockpit lantern provides natural cockpit lighting, the necessary visibility in all flight modes and during taxiing, as well as the protection of the crew from external influences. Through the glazing of the lantern, an overview of the outside space is provided:

- from the first cabin:

forward and down at zero azimuth - 16 °,

forward-up - about 160 ° and an almost circular horizontal view,

forward-side - significantly better than required by the standard;

- from the second cabin:

forward and down at zero azimuth - 6 °,

forward-side - up to 45 ° with the exception of areas shaded by the wing.

The configurations of the dashboards of both cabs are made so that they do not create shading zones outside the cab room.

In general, an overview of the outside space provided from workplaces is sufficient for the safety of flight performance and the performance of operational tasks. To improve the view of the rear hemisphere, adjustable mirrors are installed on the front and middle arches of the hinged part of the lantern. Sealing the lantern is carried out automatically when closing the locks with the help of a rubber hose laid in the under-beam profiles and temples. Compressed air to pressurize the hose is taken from the air conditioning system through the receiver. Depressurization of the cab can be done using the emergency depressurization handles. To prevent fogging of the inner surface of the glazed lamp, blowing it with heated air from the collectors of the air conditioning system is used.

Vertical plumage and stabilizer - swept in plan. The stabilizer consists of two metal structure consoles installed with a transverse angle of V = -4 °. Each console is mounted in the fuselage with two bearings and is controlled by a drive installed in the fuselage.

Kiel - metal trilonerona design. The fin of the keel is a radio-transparent fiberglass fairing with a built-in antenna of the radio station.

The rudder - a metal structure, is controlled by a drive installed in the root of the keel between the side members 2 and 3.

The integrated complex of avionics of light multi-purpose supersonic aircraft is designed to solve a number of inextricably interconnected tasks aimed at the successful completion of flight tasks.

The tasks of accurate navigation and flight control, providing supersonic and low-level flight modes with tracking of the terrain, adaptation and self-tuning of control systems are solved only using modern computer technology.

IC BREO provides an algorithmic implementation of the electronic pilot system using artificial intelligence methods and the principles of building expert systems that assist the pilot in making decisions when managing aircraft and weapons in typical situations. The system provides real-time problem solving with the capabilities of active control of the aircraft and its systems in the interests of solving the following problems:

- accumulation of data on the situation, synthesis of the situation on the basis of equally characteristic information with subsequent analysis in real time;

- choosing the optimal trajectory for flying around enemy air defense zones;

- changes and refinement of the flight route;

- a selection of recommendations on the use of tactical decisions at various stages of conducting air combat;

- classification and selection of targets (in terms of the RLPK, simultaneous detection of at least 10 targets, tracking of at least 8 targets, launching missiles for at least 2-4 targets);

- determining the number of simultaneously fired targets, the sequence and order of expendable means of destruction;

- organization of management of the operating modes of the complexes included in the IC avionics;

- the optimal use of detection tools and means of electronic suppression;

- determination of the interaction and distribution of functions between the planes of the group;

- monitoring the work of IC avionics.

IC BREO is a global information exchange system with the integration of subsystems (KSU, NK, LMS, BREK, IBKS, SU OSO) based on local multiplexed information exchange channels, which can be implemented with both wired and fiber-optic communication lines in the framework of the logical organization of combined information exchange systems.

Multifunctional means of information display and governing bodies implement the principle of "integrating information display means and governing bodies that change their function according to a certain plan in the information-control fields".

Indication support for flight and navigation modes, monitoring the operation of general aircraft equipment and combat use is implemented using three MFCIs on a liquid crystal matrix, KAI, MFPU, which are combined into a single integrated information system with a control computer with multiplex and local information exchange channels, which allows you to present an increasing amount of information from various means on a limited area of the dashboard and optimize the conditions for the perception of quantitative dashboard and naturally vnekabinnoy information to increase its visibility. The MFCI, KAI and MFPU have also been linked to the control computer center via redundant radial channels of information exchange.

To increase the survivability, two of the MFCI and the control computer are connected to the power supply system in the first category, powered by a battery and rectifier devices of an alternating current generator, main or auxiliary power units.

The principle of interchangeability of indicators allows, in case of failure of one of the three MFCIs, to provide almost the full amount of information for the performance of the flight and navigation task, and for the case of failure of two of the MFCI, a special “emergency” information frame is used, which ensures the issuance of the minimum flight and navigation information on one non-failed MFCI, necessary for safe piloting.

Integrated management system (KSU) provides:

- the required characteristics of lateral and longitudinal stability and controllability of the aircraft in the entire operational range of altitudes and speeds, angles of attack and overloads, taking into account and excluding suspensions;

- the required characteristics of stability and controllability in the mode of "refueling in the air";

- automatic limitation of limit operational values of the angle of attack (α _dop), the normal overload (n _{y dop),} the maximum speed of the instrument object (by controlling tormoznm flap) automatically limited deflection of the rudder according to the relation (δ _pH = tsV _etc., α _ist ) with manual, automatic and directorial control of the aircraft, taking into account and without taking into account suspensions for all aircraft configurations;

- disable restrictions n _{y add} from KSU remotes;

- automatic deflection of the socks of the wing, depending on the current angle of attack in accordance with a given law;

- provided manual (with the buttons "wing socks: release, cleaning") selection of the control mode of the socks; in the "release" operating mode, the installation of socks in a predetermined fixed position is ensured;

- with the “release” and “cleaning” modes switched off, the “automatic” mode is implemented in both cabins;

- Forms controlled one-time teams that implement the three-position relay control law, provides control over the position of the socks. As drives, a set of hydraulic cranes and hydraulic cylinders is used;

- automatic balancing of the aircraft by pitch and roll during cleaning - the release of wing socks, brake flap, flaps, after the suspension descends;

- flap control automatic and by crew commands and flap control control;

- brake flap control by crew commands and automatic control by instrumental speed limitation commands;

- the alarm of getting into a corkscrew and automatic withdrawal from the corkscrew - direct (flat and steep) and reverse. Provides the opportunity for training and training manual withdrawal from a corkscrew. It is realized after testing the aircraft for corkscrew characteristics and withdrawal from the corkscrew;

- reprogramming the characteristics of stability and controllability of the aircraft in order to simulate the characteristics of various types of aircraft.

KSU provides the specified priorities for controlling the flaps, unloading the RUS and pedals (for "electronic" trimmers), the choice of the control mode of the socks between the cabs.

KSU, together with a complex of on-board equipment (KBO), provides automatic and director control of the aircraft:

- automatic stabilization of the angular position by pitch, roll (course),

- automatic stabilization of barometric altitude in the entire permitted range of heights and speeds;

- in the automatic height stabilization mode, ensure stabilization of the zero value of the angle of inclination of the trajectory;

- automatic and director stabilization of barometric altitude in the operating modes "Landing", "Return", "Route";

- automatic stabilization of instrument speed (through the stabilizer channel), automatic control of the aircraft in horizontal and vertical planes in the Route mode, automatic and director control of the aircraft according to signals from the BCVS, which is part of the BWC, when flying along the flight path, return and pre-landing maneuver in horizontal and vertical planes with access to the coverage area of ground landing equipment;

- combined aircraft control mode;

- command control of engine thrust by means of recommendations through the ARO and an indication system of the required engine operating mode in the "low gas - maximum" range in order to prevent the aircraft from leaving the safe flight speed ranges for the current aircraft configuration;

- automatic and director control of the aircraft during approach, landing mode. Automatic and director control of the aircraft during the second approach (according to commands and signals from the BWC, the "Repeated approach" mode);

- automatic reduction to horizontal flight at the command of the crew and from the automatic modes "Departure" and "Automatic withdrawal from a corkscrew" after their implementation;

- automatic withdrawal from collisions with the ground and from fragments of their own ammunition;

- automatic stabilization of geometric height (low altitude flight);

- automatic and director control of the aircraft according to signals from the BCVS KBO in horizontal and vertical planes in the modes of "Guidance", "Attack".

KSU provides the formation of redundant information on altitude and speed parameters in accordance with the requirements of aircraft systems, including display and registration systems.

KSU sensors provide the formation of information about the primary altitude-speed parameters through four channels:

Total pressure from 9 to 2830 (± 0.3 ... 3.1) mm Hg;

Static pressure is from 9 to 807 (± 0.3) mm Hg;

Braking temperature from -60 to +350 (± 0.1) ° С;

Local slip angle from + 30 ° to -30 ° (± 0.24 °);

Local angle of attack from -25 ° to + 75 ° (± 0.24 °).

Based on these parameters, the KSU calculators, taking into account aerodynamic corrections, calculate the current values of the true air speed (V _East ), instrument speed (V _CR ), number M, absolute (N _abs ) and relative (N _rel ) barometric altitude, as well as true values angles of attack and slip.

In addition to measuring the aerodynamic angles α and β with the help of DAEs, the KSU provides for the calculation of the values of these angles and their derivatives using all the information available and embedded in the KSU. This information is used in the KSU to improve the characteristics of the DAU signals to control mechanical failures of the DAU and to expand the range of measurement of the angle of attack to 70 °.

KSU provides the solution to the problems of the system of restrictive signals and the formation of:

- redundant signals α _add , n _{y add} , Vmin, Vmax, Mmax for various aircraft configurations and various suspensions with issuing them for display and registration;

- sound alarms and visual indication: about approaching n _{y add} , when the overload limiter is off and in configurations with disabled or missing ODA; to Vmin, Vmax to a dangerous altitude, n _{y extra} during manual piloting in the "landing" mode.

Formation of reserve signals of pitch and roll angles, ensuring safe piloting of the aircraft when returning to the base airfield after a failure of the source of angle signals and issuing them to the display and registration systems.

The system provides control of the health of its own equipment, control of input information from the systems (by signals of their health) and sensors, including their heating, shutdown of failed equipment, systems, sensors, control modes with the issuance of relevant information in the display, alarm, registration and ARC systems.

To ensure proper functioning of their own equipment in KSU, the following built-in control methods are used: "inter-channel selection" (quorumation), "controlled pair" - comparison of signals of two identical devices, "channel-model" - comparison of device signals and its models, software and hardware control. The combination of built-in control and backup management tools used provides standardized reliability indicators of KSU.

An integrated control system provides the deviation of the controllable surfaces of the aircraft (downstream) at angles:

ailerons: up - 1α, I = 35 °, down - 1α, I = 30 °;

stabilizer δ = + 15 ° --40 °;

rudder δрн = 0 ± 30 °;

wing socks = δnk = 0-30 °;

flaps δzak = 0 - removed; 10-20 - take-off, 20-40 - landing;

brake flap δтщ = 60 °.

Permissible control errors - a larger value of ± 1 ° or 10% of the measured value, within the limits of the available forces of the drives.

KSU provides ailerons float and hovering within the specified deviations: _{Δδ ezav} , _{Δδ evspl} .

The architecture and construction of the computing process of the KSU provide the formation of identical control signals from the outputs of the reserves of the digital computer to control the spools of electro-hydraulic drives.

The construction of the control spool control circuits of the drives provides synchronization of the position of the independent spools of both chambers.

For relay control of socks, four electrohydrocranes are used (2 hydraulic systems, 2 released positions).

With the sock control modes turned off (socks: cleaning, bleeding) in the rear cab, the priority in choosing the sock control mode belongs to the front cab, otherwise, to the rear cab. Priority change is signaled in the front cab.

By commands from the control panels, the KSU generates signals for the release (cleaning) of the flaps to the “removed”, “landing” position; if there are no specified commands from the KSU control panels, the flaps are released to the “take-off” position and are automatically retracted in accordance with the position of the chassis and the reserve Vmax (with flaps extended), KSU monitors the position and timing of the release (harvesting). The specific values of the flap deflection angles in the take-off and landing positions are quick-changeable.

Allowable flap mismatch angle is 5 °. If this discrepancy is exceeded, the KSU provides the removal of control signals for flap exhaust release and shutdown and shutdown of the flap drive with an alarm about the failure (asymmetry) of flap control and the generation (if necessary) of signals for the flap brake.

If the flap position commands on the KSU console of the rear cockpit are disabled, the priority for flap control belongs to the front cockpit. If one of the two flap position commands is available from the KSU control panel of the rear cockpit, control priority is transferred to this cockpit. A change of priority is indicated in the front cab.

KSU generates one-time teams that implement the relay control law (cleaning - release) by the brake flap, according to crew commands from both cabs or by a signal of restriction V _ave . As a drive, a hydraulic cylinder and a hydraulic crane are used.

If there are no control commands from the RUS unloading controls and pedals of the rear cockpit, the KSU provides priority for unloading control from the front cockpit, otherwise from the rear cockpit.

KSU provides information about the deviation angles (position) of the controlled surfaces in the aircraft display system.

In the “reprogramming” mode of operation of the control system, the control gradients are changed by changing the dependencies according to the signals from the redundant three-position switch installed in the control panel compartment or in the rear cabin and protected from accidental switching. Turning the mode on and off is carried out with the MFP CCD.

In the event of a failure of three receivers of full or static pressure, the control panel switches to internal correction signals. In this case, restriction of operating conditions is allowed. KSU provides a manual transition to a backup version of open source software. The transition is provided with the KSU POOL by controls protected against accidental switching on.

KSU provides an automatic transition to the required control algorithms when the “Corkscrew” command appears (it is implemented when the mode automatically withdraws from the corkscrew).

The KSU will automatically switch to the backup version of the STR after detecting a failure of the control version of the STR.

To ensure KSU adaptability to the characteristics of a family of light aircraft, control laws provide:

- increasing the adaptive capabilities of control laws in order to provide the required stability and controllability characteristics in aircraft configurations associated with the use of external suspensions, including asymmetric options, in extended ranges of weight, centering, overloads and speeds;

- reduction in pilot loading under the action of asymmetric roll and / or yaw moments, the appearance and instability (from plane to plane) of which are likely at angles of attack greater than 15 °, when flying in conditions of partial or complete breakdown of the wing flow;

- a significant improvement in the coordination of roll and yaw movements in order to maintain the sliding angle in the area in which it is possible to provide acceptable stability and controllability characteristics when controlling from both roll RUS and pedals.

The main features of the laws of control of the longitudinal channel of KSU:

- the use of feedbacks on the pitch angular velocities ωυ, the angle of attack ωα, the angle of attack signal α and the normal overload signal n _y (the angle of attack signal is used in the entire field of application of the aircraft to compensate for nonlinearities of aerodynamic characteristics and in a limited range of flight modes, as the coordinates of aircraft control , at the same time, at high speed heads, the signal n _y ) is used as the control coordinate;

- introduction to the number of signals for automatic shockless switching of the "Takeoff-Landing" and "Flight" operating modes, in addition to the signals of the position of the flaps of the one-time commands "landing gear removed" (released);

- introduction of signals to the structure of the CDS for automatic rebalancing of the aircraft during the transition from the flight configuration to the take-off and landing and back when controlling socks and the brake flap;

- introduction to the structure of the CDS of a nonlinear prefilter that limits the magnitude of the slew rate of the control signal when moving the control knob;

- the transition in the flight configuration of the aircraft to the integral control law, applied only to limit n _y .

The main features of the laws of control of the side channels of the KSU:

- the use of feedback signals as signals to ensure stability, acceptable dynamics and controllability of an aircraft with a CDS of angular roll speeds, slip angle and rate of change of the slip angle;

- there is a cross connection from the control stick to the direction channel to ensure the required roll speeds at large angles of attack and "rough" coordination of roll and yaw movements;

- to improve coordination, the use of a signal of the rate of change of the angle of slip β

- a limitation of the rudder deflection is provided from the conditions of structural strength, but with the possibility of maximizing the use of rudder deflection at large angles of attack.

To reduce the complexity of piloting when parrying the external moments of the roll or yaw and to improve coordination of the movements of the roll and yaw, the use of integral control laws is provided both in the aileron channel and in the direction channel.

The angle of attack is limited in accordance with the value of α _extra generated by the KSU for all aircraft configurations. Values of α _add and n _{y add} correspond to the maximum deviation of pitch pitch.

The power supply of the KSU is carried out from two independent buses, while ensuring that there is no interruption in the power supply of the KSU. KSU is supplied with power from a power supply system consisting of two independent AC channels (sides) with a nominal voltage of 115/200 V, a nominal frequency of 400 Hz and two independent secondary buses operating from DC rectifier devices with a nominal voltage of 27 V. GOST 19705-89.

KSU power supply is carried out according to the following circuits:

alternating current - heating DAP, DAU and LDPE;

DC - power supply to the computing units of the system.

When the DC voltage drops below 16 V, the power supply units of the KSU are turned off. Recovering voltage above 16 V is equivalent to applying power. When starting the APU, the DC bus disconnects. A simultaneous break and simultaneous voltage drops across two buses are excluded.

KSU for direct current is a consumer of category 1, and for alternating current: heating DAP, DAU, LDPE and fan power (if necessary) - a consumer of 2 categories.

Electric power consumed by KSU, no more than:

alternating current - 1400 VA;

DC - 900 watts.

KSU readiness time after power-up is no more than 3 minutes, taking into account the time of ground control. There is an interruption and blocking of ground control from the BWC remotes.

KSU executive drives are electro-hydraulic steering drives (type RPD-15A in the roll channel and directions, type RPD-17A in the pitch channel).

Electro-hydraulic drives are powered by two independent hydraulic systems. The operation of the drives (type RPD-15V) with effort is ensured in the presence of working pressure in both hydraulic systems. In the absence of pressure in one of the hydraulic systems, steering surfaces are controlled from another hydraulic system, while the force developed by the actuators is approximately halved. In the pitch channel, the required efforts are provided.

The flaps control system ensures their movement and fixation in the released and retracted positions. The control system is powered by a common hydraulic system.

The release and cleaning of the flaps is controlled from the control panels and from the KSU computer, the signal of which is fed to the flaps steering gear; through flexible shafts, the rotational movement is transmitted to rotary gears, which directly controls the flaps. Each of the flaps is connected to two rotary gears. On each flap, redundant position sensors are installed, the signals from which are fed to KSU computers. When the flaps reach the set position, the control signal is removed. In a predetermined position, the flap is held by self-braking of rotary gears and a brake. The angle of deviation of the flaps to the take-off position is 10-20 °. The flap deflection angle to the landing position is 20–40 °.

The position of the flaps is controlled automatically and mnemo frame by the MFCI crew.

KSU controls the synchronization of the flap exhaust release. When the angle of the mismatch of the flaps is more than 3-5 °, the control panel removes the control signal, the flaps stop, the crew receives a message on the MFCI "flaps-failure".

KSU implements the automatic control flaps control mode. On take-off after landing gear, the system starts flaps cleaning, on landing, after landing gear flaps are released to take-off position.

The direction of rotation of the shaft is reversible, depending on the supply of an electric signal to electromagnetic cranes. The response time of the drive from the moment of turning on (turning off) the electromagnetic crane and the turning-on valve to the start of rotation (stop) of the drive shaft is not more than 0.15 s.

KSU manages wing socks automatically depending on the current angle of attack. Sock deflection angle 25 ° and 30 °. With 2 ° flaps extended, the socks automatically deflect at a 25 ° angle.

Each wing console has two sections of socks: root and end.

Each section is controlled by a booster and a common hydraulic system. In a given position, the wing socks are held by the pressure of the working fluid. To ensure control control and compensation of asymmetries in the control of the toes of the wing, redundant position sensors are installed on each section, the signals from which are fed to KSU calculators. In the event of a control failure of at least one of the sections for any reason, the system generates a failure signal.

KSU interacts with the following equipment:

- On-board digital computer system BTsVS;

- A comprehensive system of electronic indication and control KSEIU;

- strapdown inertial navigation system BINS;

- low-altitude radio altimeter (type RVM-95);

- radio system of near navigation and landing (VIM-95, DME);

- satellite navigation system (A-737-003);

- weapons management system (SLA);

- aircraft hydraulic system with command sensors about the presence of pressure;

- aircraft power supply system (SES);

- airborne registration device ("Carat", ZBN, EBN);

- system of onboard measurements ("regatta");

- remote engine control system;

- Aircraft equipment management system;

- chassis status sensors: “front left compressed” - single-channel, power from the left side, “right front compressed” - single-channel, powered from the right side, “chassis removed” - four-channel, powered from four KSU reserves on the front landing gear, command " cleaning (release) chassis released "(single-channel);

- aircraft intercom equipment (MSPU);

- Hydrococks to control wing socks, brake flap;

- buttons to turn off automatic control and bring to the horizon on the RUS;

- handle unloading mechanism in the pitch channel (type MP-10) and controls for unloading the RUS and pedals;

- socks control mode selection switches (in two cabs);

- Switches for switching on the "Leave" modes;

- switch of options of the "Reprogramming" mode.

Information exchange between interacting equipment and KSU is carried out in accordance with GOST 26765.52-87. In this case, the OA terminal devices for communication with the ICIE are part of the KSU.

The control system for general aircraft equipment provides:

- reception and digitalization of 110 analog and 330 discrete signals from interfaced systems and sensors of general aircraft equipment;

- the issuance of 92 control commands to actuators and actuators, as well as 48 control signals for flashers;

- receiving information on 16 code communication lines with parameters meeting the requirements of GOST 18977-79 and RTM 1495-75, with an information transfer rate of 50 kbit / s + 1%;

- information transfer via 8 code communication lines with parameters meeting the requirements of GOST 18977-79 and RTM 1495-75, with an information transfer rate of 50 kbit / s + 1%;

- reception and transmission of information to the ICIE 4 information exchange channel (primary and backup).

The system provides control functions for general aircraft equipment in manual and automatic control mode (if this is provided for by the controls and elements of each system separately). The manual control mode is carried out by receiving a command to execute the control sequence diagram of the aircraft system actuator from the controls located in the cockpit. In the automatic control mode, the signs of activation of the CCA control cyclograms are generated on the basis of information received by the system from aircraft systems. Manual control has higher priority than automatic control.

Control of aircraft equipment by the CCA OS system is carried out by comparing the generated control commands with the information of the received systems as a result of its implementation and using the “request” commands initiating the launch of the built-in self-monitoring of avionics systems. Information on the results of the control of the CC OS is transmitted to the registration system (objective control) and the digital computer. SU CCA provides the issuance to the registration and control system (objective control system) through the ICIE information exchange channel of values of current parameters, event codes and codes of failed system blocks identified at all stages of the facility.

The system, in case of a failure of the digital computer, functions as the controller according to the MKIO, and in the event of a failure of the MKIO, the OSO control unit is connected to multifunction indicators (MFCI) and to the OSO control panels by radial communications.

The system is powered from power supply systems (SES), providing two secondary channels (sides) of direct current with a nominal voltage of 27 V, the quality of electricity in accordance with GOST 19705-89 for systems with rectifier devices.

In this case, alternating current (115 V / 400 Hz) can come both from the main generators and from the auxiliary power unit (APU) generator, and in case of failure of the SES of alternating current, the OS CCS is powered from the battery.

The BPV conversion and calculation units are connected to two SES channels and operate normally with power quality parameters for one channel - for normal and emergency modes (in steady and transient modes), while the second power channel can be powered with electricity with quality parameters for emergency mode.

With a simultaneous interruption in the power supply through two channels, the system does not fail and does not give false signals and commands (except for their removal), when restoring the quality of electricity in at least one power supply channel of the SES to a normal or emergency value, the system is automatically restored in 1 second without crew intervention.

The integrated power supply system (SES) is designed to provide electric power to the aircraft's on-board systems (power receivers) in all the stipulated modes of its operation. The power quality parameters produced by the system comply with GOST 19705-89 and MILSTD704E.

SES consists of two subsystems:

SES ~ - AC power subsystems with a rated voltage of 115/200 V at a stable rated frequency of 400 Hz;

SES = - DC power supply subsystem with a rated voltage of 27 V.

SES ~ and SES = include alternating and direct current generation systems.

The work of the SES is as follows.

At the aerodrome, with non-running marching engines, alternating current power is supplied from the aerodrome source through the connection unit or from the APU generator - in stand-alone mode. Joint operation of the airfield source and the APU generator is possible, while the left side bus is fed from the airfield source, and the right side bus is fed from the generator of the APU.

After starting the main power plant, drive generators are connected - the main sources. The scheme is made with the priority of the main sources, therefore, even when the airfield source and the APU generator are working, the left side bus is powered from the main left drive-generator, and the right side bus from the right drive-generator.

In flight, in case of failure of one of the drive generators or when it is turned off (due to a failure (stop) of the engine of the main power plant), the APU is started and the APU generator is automatically turned on, which is connected to the bus of the inoperative main generator (the pilot turns on the switch of the APU generator in order to when touching the runway, when landing, the generator did not turn off).

The power of each of the drive generators and the APU generator - 16 kVA (up to 10 km) allows you to provide electricity to all on-board systems of the aircraft, even when one of these sources is in operation.

During the operation of any AC power source, the left and right side tires are automatically combined and powered from the specified source.

Depending on the number of working sources, the power addresses of the left and right side buses are changed by means of switching contactors.

In the presence of alternating current on the left and right side tires both WUs work. The left VU feeds the left side of the direct current (bus lev. Avar. And the bus left. Off), the right VU feeds the right side of the DC (bus right. Avar. And bus right. Off).

Each of the control units works in parallel with one battery left and right, respectively.

In the event of a failure of one of the VU, the tires ("bead") are combined by the contactor of the on-board association and powered by one operating VU. The power of each of the VU - 6 kW is sufficient to provide power to all on-board systems.

When both WUs are disconnected due to de-energization of both busbars ("boards") of the alternating current, disconnected lion buses are disconnected. off and right. off, and the tires are lion. avar. and right. Avar. are combined and flooded by both batteries.

When starting the APU (connecting the electric starter of the APU) the left and right side tires:

in flight - are disconnected;

on the ground - they are disconnected when starting the APU from one side;

on the ground - are combined when starting the APU from two sides.

The electronic indication complex provides flight and navigation modes, monitoring the operation of aircraft equipment and combat use with the help of three MFCI, KAI, MFPU, which are combined into a single integrated information system with control systems of a computer, multiplex and local channels of information exchange, which allows not only to present an increasing volume information from various means on a limited area of the dashboard, but also optimize the conditions for the perception of quantitative instrumentation and natural out-of-cabin information, to increase its visibility. The MFCI, KAI and MFPU communicate with the controllers of the computer center also through redundant radial channels of information exchange.

In order to increase the survivability of the MFCI, the controllers of the computer are connected to the power supply system in the first category, using an alternating current generator of the main or auxiliary power units and batteries.

The principle of interchangeability of indicators allows for the failure of one of the MFCI to provide almost the full amount of information for the flight and navigation and combat missions, and for the case of failure of two of the MFCs, a special emergency information frame is used to provide the minimum flight and navigation information on one non-failed MFCI necessary for safe piloting.

The logic of the distribution of indicator information on the screens of the MFCI and KAI is based on a number of provisions.

Navigation mode includes:

- take off;

- flight along the route;

- "guidance" and "attack";

- return to the landing airport;

- landing;

- re-approach.

The mode of combat use includes:

- guidance;

- attack;

- exit from the attack.

Information on indicators (with a 4-indicator version of the cab - 3 MFTSI + KAI) is distributed as follows.

The basic information for flight and combat use is placed on the KAI - the position of the aircraft in space, data on altitude, speed, direction of flight, data on the target, area of destruction and instructions on the use of weapons. Also, warnings about dangerous flight modes and failures on board can be issued.

The left MFCI provides flight and navigation information in an amount sufficient to perform tasks in difficult weather conditions (clouds, above the sea, out of sight of the horizon, at night, etc.).

The average MFCI represents a tactical situation, navigation data or cartographic image.

The right MFCI provides the pilot with information on the operation of aircraft systems, engines, equipment, issues recommendations for actions in special cases, in case of erroneous actions by the pilot during the flight stages within the framework of the expert system, and information from television-optical weapons.

In case of failure of any of the MFCIs, information from the failed indicator can be installed on any other MFCI.

The multifunctional MFPU control panel can be located in two cabins of the aircraft: in the first cockpit - on the KAI, in the second cockpit - on the dashboard and is designed to perform the following functions:

- management of the complex (in conjunction with the MFCI);

- control of the operating modes of the complex and systems interfaced with the complex (switching of the operating modes of the complex, the operational change of the flight task parameters, the use of PUIs as control panels of the complex systems and systems interacting with the complex);

- Indication and editing of individual flight task parameters. Management of the operating modes of the complex is carried out using the main pages of management, management of the modes of operation of systems - using the pages of system management.

A television system for the objective control of combat training is installed on board the aircraft, the T-SOK-UBD is a set of devices for monitoring, recording and displaying MFCI video information captured by television cameras during the flight, as well as for recording audio conversations between the crew and with the control point.

The T-SOK-UBD system is designed to provide:

- video surveillance of airspace in the front hemisphere in the direction of movement of the object;

- video surveillance of the results of the combat use of TSA;

- registration of video information from 4 television cameras and from 3 MFCI to an easily removable solid-state drive;

- registration of sound negotiations of the crew;

- issuing video signals from one of the television cameras to the BCCO to display the image on the MFCI.

Television information (TV) is transmitted to the KAI from a color video camera and from television-optical lesion systems and through the switching unit of television signals BKTS can be quickly output to the MFCI.

A video recorder is installed on board that records the indication from the KAI with the cockpit space and radio exchange, and if necessary, the information from the KAI can be copied to any MFCI.

The main task of the KAI is to issue targeted information to the pilot against the flight. Information "flexible" - varies by stages of flight. Stages: take-off, flight en route, combat use, approach. The basis of the flight information is the velocity vector, which receives data characterizing the movement of the aircraft from the inertial system, and when it is rejected from the sensors of the angle of attack and slip.

When the "Take-off" mode is executed, the following symbols are displayed on the format:

- aircraft circuit;

- roll scale (fixed);

- pitch scale (movable);

- course scale;

- digital counter of the current height;

- digital instrument speed counter;

- mode indicator - the inscription "VZL" (take-off).

When the "Route" mode is executed, the following symbols are displayed on the format:

- aircraft circuit;

- roll scale (fixed);

- pitch scale (movable);

- course scale;

- control mode indicator: "DIR" - director, "AVT" - automatic;

- the velocity vector, receiving data characterizing the movement of the aircraft from the inertial system;

- the symbol "danger zone" - appears on the screen 4-5 seconds before reaching the height from which it is necessary to begin the maneuver to move from a dangerous height;

- the symbol "OTV" (lapel to the ground) is displayed 2 seconds before reaching a dangerous altitude and is removed when entering safe flight conditions;

- digital height meter in meters;

- digital counter for instrument speed in km / h;

- movement mode in the vertical channel: "HIGH" - stabilization of height, "PG" - bringing to the horizon, "STB" - stabilization of angles, "DRIVING" - mode of moving away from the earth;

- digital counter of the distance to the navigation point.

In the “RETURN” mode (return to the landing aerodrome), a frame identical to the frame of the “ROUTE” mode is formed on the KAI with the following difference: instead of specifying the “ROUTE” mode, the text “EXPLOSION” (return) is formed.

In Landing mode, the following symbols are displayed on the format:

- aircraft circuit;

- roll scale (fixed);

- pitch scale (movable);

- course scale;

- digital counter of the current height;

- digital instrument speed counter;

- set course - runway course;

- flight mode - "PIC" (landing);

- control mode indicator: "DIR" - director, "AVT" - automatic;

- the velocity vector, receiving data characterizing the movement of the aircraft from the inertial system;

- synthesized "pseudo" runway, matching in location and angular size with the real visible;

- range to the end of the runway;

- a symbol indicating a deviation from the course and glide path beacons. The KAI mnemo frame for RETRIEVE mode corresponds to the mnemo frame for ROUTE mode with the following differences:

- the mode indicator is indicated - "PZ" (re-entry);

- the set course is formed in accordance with the reentry path.

The KAI mnemo frame corresponds to the “ROUTE” mnemo frame with the symbols “Danger zone”, “Turn away from the ground”, and the text symbol “DRIVING” appears.

The nomenclature of the flight and navigation mnemonic frames and mnemonic frames for monitoring the state of aircraft systems on the MFCI indicators is the same and includes the following mnemonic frames:

- combined flight and navigation information: IKP + NPP;

- navigation and planning information: "REVIEW", "ROUTE", "ITO", "LANDING";

- control of the state of aircraft systems: formats - "TYPICAL (main)", "ENGINE", "HYDROPNEUMO", "FUEL", "SJO", "SDU", "SES", "PREPARATION".

A specific mnemonic is selected by the pilot using the multifunction buttons framed by the MFCI. The menu is present on all mnemo frames, at the top of the screen of the right MFCI.

The left MFCI is used as a flight-navigation indicator and carries information sufficient to perform any flight maneuver, as well as to perform aircraft navigation to the selected navigation point.

On the frame in the format of the flight and navigation information indicates the position of the aircraft relative to the horizontal plane. Horizontal and vertical speeds, overload, angle of attack, velocity vector, barometric and radio heights, set and dangerous flight heights, heading scale, set and current courses, distance and angular deviation from the current point to which the flight is carried out, deviation from course and glide path are displayed beacons (in the LANDING mode). In addition, barometric pressure, time and fuel reserve are displayed.

The central MFCI is used as a flight-planning indicator and carries information about the position of the aircraft relative to a given path in the horizontal plane during take-off, flight along the route and approach (planned flight picture, or an electronic map of the area).

The right MFCI is used as an indicator of monitoring the state of aircraft systems and carries information about the configuration of the aircraft, the operation of all systems and assemblies. The main form of presentation of information is mnemonic. The depth of control is sufficient for analysis and decision-making on the operation of systems by both flight and technical personnel without additional connection of control panels. If necessary, information from television-optical means of destruction is displayed.

The scheme of the control system of suspended objects when using a universal device on an aircraft operates as follows. The control signals from the weapon control system are received by all devices for attaching external suspension objects via a single communication, with branching in accordance with the logical numbers of the objects, which allows the commissioning of a specific unit responsible for starting (disconnecting) the corresponding suspension object.

The use of a universal device for mounting external suspended objects (a universal aviation pylon) on an aircraft will increase the payload, improve the aerodynamic qualities of the wing, reduce movement resistance, reduce weight, improve flight tactical characteristics and, in general, increase the efficiency of the intended use of the aircraft .

Placing the blocks in the pylons can significantly simplify the process of controlling suspended objects and reduce the weight of the aircraft, as well as reduce preparation time for flight when changing the type of suspended object, because to control another object, just replace the block. The block is implemented as a separate module, and its replacement is extremely simplified.

Autonomous on-board oxygen system has the following advantages:

- reducing the time of preparation of the aircraft for flight;

- increase the operational manufacturability of aircraft maintenance;

- Integration of the aircraft airborne oxygen system with the control system for general aircraft equipment (SU OSO) and the multifunction display system - MFCI:

- issue and indication of the modes "EXIT ON THE MODE", "OXYGEN FAILURE", "OXYGEN NORM"; "CURRENT VALUE OF THE PARTIAL PRESSURE OF OXYGEN"; "INDICATION OF RESPIRATORY PHASES";

- lack of need for ground-based facilities that produce and transport oxygen;

- the removal of restrictions on the flight time of the aircraft associated with the limitation of the oxygen supply on board, which is especially true when flying with refueling in the air.

Depending on the flight conditions, the airborne oxygen system provides:

- production using an onboard oxygen production unit (for example, BKDU-130) and breathing with the required partial pressure of oxygen;

- emergency supply of oxygen from the oxygen system of the ejection seat in case of failure of the onboard oxygen production unit;

- the supply of oxygen-enriched gas medium under excess pressure to the breathing line and the supply of compressed air to the chambers of the anti-overload device, depending on the magnitude of the overload;

- oxygen supply to the oxygen mask during bailout, subsequent descent in the chair and landing (splashdown);

- protection of respiratory organs from substances harmful to breathing.

The oxygen devices included in the system (for example, type KP-130) provide pulmonary-automatic supply of the respiratory mixture into oxygen masks. In the event of a valve mechanism failure in the pulmonary machine, manual activation of the continuous supply of the respiratory mixture into the mask (with the handle on the gearbox) is provided. In the absence of pressure at the inlet of the manual gearbox, the pilot's breathing with ambient air from the cockpit is provided through an oxygen device suction valve.

The pressure regulator (for example, type RD-130) is designed to create excess pressure under the mask during overloads in order to increase the crew's tolerance of overloads. It is controlled by an automatic pressure controller (type AD-16V), which simultaneously provides for overloading of an overload suit (PPK) during overloads.

The oxygen system of the ejection seat (type KSKK-130) is an emergency source of oxygen to provide oxygen to crew members:

- during depressurization of a cabin with a decrease in its barometric pressure to a level corresponding to a "height" in the cabin of more than kilometers;

- in case of failure of the control panel;

- during bailout followed by descent by parachute.

To prevent the flow of oxygen from the included KSCS to the oxygen device of another crew member, a check valve is used. The gas supply sensor for breathing (DPGD) is designed to detect the presence of a flow rate of oxygen-enriched air from the BKDU into oxygen masks for the breathing of the crew.

The work of the aircraft airborne oxygen system is as follows.

The power supply BKDU (Fig) is performed after starting the auxiliary power unit (APU) switch "KS - on-off" of the common bracket for turning on all systems. In this case, the gas analyzer is warmed up and adjusted for 7.5 minutes. On the screens of the MFCI in the cockpit, an indication of the “exit to mode” signal should appear in the form of a flashing symbol, for example, blue and white.

Within 0.5-1 min after the start of all the engines of the main power plant of the aircraft, the accumulated condensate, contaminants and insects are discharged from the air intake lines from the engines and the APU. Then, according to a signal from the on-board control system, the crane automatically switches to the water discharge unit from the drainage position to the compressed air supply to the air control unit and to the pressure control devices, through a special filter that cleans the air of dust, aerosols (including toxic ones) and drip moisture . Drip moisture is collected by the filter and discharged overboard. The air purified by the filter enters the BKDU inlet and to the pressure switches.

Inside the CCU, the air through the gearbox enters the distribution valve block and then to the adsorbers, where it is divided into oxygen-rich DHA and nitrogen-enriched mixture.

The respiratory mixture from BKDU comes through DPGD sensors and oxygen devices KP in oxygen masks of pilots. The nitrogen enriched mixture is dumped overboard.

The partial pressure of oxygen in the DGS is controlled by a gas analyzer included in the BKDU, which is connected to the cabin by an equalization line. The control of the excess gas pressure at the BKDU outlet is carried out by the pressure signaling device included in it, the static cavity of which is also connected to the equalization line. In order to prevent air flow from the cabin to the control cabinet during pressure testing of the cabin, as well as to protect it from contaminants entering the equalization line, a fitting with a non-return valve is installed on its end part.

Information about the state of the gas generator from the gas analyzer and from the pressure signaling device enters the electronic control unit BKDU, which issues the following signals to the on-board display system:

- "exit to the mode" within 7.5 minutes from the moment the power supply BKDU;

- "oxygen is the norm" - at a partial pressure of oxygen in the gas-pressure generator of at least 25 kPa on the ground, at least 15 kPa in flight and at an excess pressure of the gas-pressure generator (relative to the pressure in the cabin) of more than 20 kPa;

- "oxygen - failure" - at a partial pressure of oxygen on the earth of less than 25 kPa, in flight less than 15 kPa (up to 13 kPa) for 10 min or less than 13 kPa for 30 s, with an overpressure pressure of the gas generator (relative to the pressure in the cabin) less than 20 kPa for 1 min;

- "current value of the partial pressure of oxygen" - the current output signal of the gas analyzer.

In order to distinguish between the ground and flight conditions in the airborne control unit from the on-board system by the electronic unit, a ground signal should be issued.

The quantitative characteristics of the signals can be as follows:

Parameters of compressed air at the system inlet.

At the entrance to the BKDU:

- overpressure relative to atmospheric pressure at a flight altitude of 100 to 1000 kPa (1 to 10 kgf / cm);

- temperature from zero to + 95 ° С (short-term up to 40 ° С);

At the entrance to the blood pressure during overloads of more than 2 units:

- overpressure relative to atmospheric pressure at a flight height of from 150 to 1000 kPa (from 1.5 to 10 kgf / cm);

- temperature from zero to + 80 ° С (short-term from minus 40 to + 95 ° С).

Parameters of the produced BKDU of the respiratory gas medium:

- the oxygen content to the "height in the cabin" 8000 m (according to OST 1 00831-77);

- performance provides breathing of two crew members with pulmonary ventilation up to 30 l / min each.

An electric command to automatically turn on the KSCK is issued by the on-board system when an oxygen-failure signal is received at a "height in the cabin of more than 8000 m. At a" height in the cabin "of more than 10 km, the KSCS is automatically turned on without waiting for the oxygen-failure signal.

Visual indication to the pilot: "gas supply for breathing" - light signal oscillating in time with the phases of respiration; "exit to the mode" - a light signal; "oxygen-norm" - a light signal; "oxygen failure" - light signal; "emergency oxygen" - light signal and voice warning; "partial pressure of oxygen".

Parameters recorded on the on-board drive: "air supply to the BKDU is on"; "air supply to BKDU is disconnected"; "BKDU output to the mode"; "oxygen is the norm"; oxygen failure; "emergency oxygen"; "gas supply for breathing"; the current value of the partial pressure of oxygen from 5 to 25 kPa.

Indication form:

"exit to the mode" - a blinking mnemonic sign symbolizing the oxygen system;

"oxygen is the norm" - the same sign glows constantly:

"oxygen - failure" - a change in the color of the sign to red, accompanied by the text information "oxygen - failure".

The indication of "exit to the mode" should remain until the crane switches to the air handling unit from the drainage position to the compressed air supply position in the control panel. If the time to this point exceeds 7.5 minutes from the moment the power supply of the BKDU is turned on, the indication "oxygen - normal" and "oxygen - failure" will be blocked during this period. If the signal “oxygen - norm" or "oxygen - refusal" does not arrive from the control panel for 5 seconds after entering the mode, the on-board control system generates information "oxygen - refusal" at the MFCI.

The signal about the presence of the supply of DGS to oxygen devices comes from the DPGD sensor. During normal operation of the “pulmonary automaton” of the KP oxygen device and the DGS flow through it, corresponding to pulmonary ventilation of more than 7.5 l / min, the DPGD sensor gives a signal that varies in phases of respiration, which is sensed and displayed by the on-board display system. When the pulmonary machine fails in the open position or when the continuous supply (using the handle of the oxygen device) is turned on, the DPGD sensor gives a continuous signal of maximum level. In case of failure of the pulmonary automaton in the closed position, or in the absence of supply of DGS from the BKDU, or when the DGS flow corresponding to pulmonary ventilation is less than 7.5 l / min, the DPGD sensor produces a continuous signal of a minimum level. The method of visual display of this signal is determined by the design of the on-board display system.

The form of indication is a blinker-type image oscillating in time with the phases of breathing against the background of a mnemonic sign symbolizing the oxygen system.

The oxygen system of the ejection seat KSCK provides both manual and automatic inclusion. The electric command for automatic inclusion is issued on KSKK by an onboard control system:

- in the case of depressurization of the cabin with a decrease in the barometric pressure in it to a level corresponding to the "height" in the cabin of more than 8 km and receiving an oxygen-failure signal from the control panel;

- at a "height" in the cab of more than 10 km, an electric command to turn on the KSCK is issued without waiting for the signal "oxygen - failure".

Simultaneously with switching on the KSCK, the on-board automation issues an alarm oxygen signal to the display system.

The display form of this signal is text and voice alerts.

To prevent the flow of oxygen from the included KSCS to the oxygen device of another crew member, a check valve is used.

When ejecting, the oxygen supply from the BSCC is turned on by a mechanical device.

After the flight is completed and the engines are stopped, the on-board control system provides reverse switching of the condensate drain valve of the UVP device to the drainage position, while eliminating the possibility of de-energizing the electromechanism of the crane before switching it to the drainage position.

Claims (3)

1. A light supersonic multipurpose aircraft containing a fuselage with a cockpit in the bow, a brake flap in its upper part, side air intakes bounded by a wing, landing gear, the main power plant with two turbojet engines, a swept wing with a front root influx of a lively shape, mechanization and deflectable socks with ledges of the leading edge, single-tail vertical tail and horizontal tail, characterized in that the root influx of the wing is made with a span of l _n not less than 0.15 and not more than 0.18 r the wing span l _{k of} the wing console from the base of the influx and with the chord b _{n of} at least 0.55 and no more than 0.6 of the local chord b _{k of the} wing, the ledge of the leading edge of the deflected wing tip is made at a distance of 0.55-0.65 of the magnitude of the span l _k the wing console from the base of the influx, with a chord b _{n of} at least 0.075 of the local chord b _{k of the} wing, without taking into account the ledge, the wing has a reverse sweep of the trailing edge of the wing and is made with universal devices for mounting external hanging objects placed symmetrically on each console horizontal plumage made in the form of differential-controlled stabilizers, the axis of rotation of which intersects with the axis of the engines of the main power plant, and the trailing edge protrudes from the end of the nozzle units, turbojet engines are made ultrasonic, placed in the rear of the fuselage with the horizontal part of the tail unit located between them, auxiliary power unit located at the bottom of the middle part of the fuselage behind the cockpit, the inlet openings of the air intakes are located at the bottom and are limited to the fuselage, in The upper part of the air intake, limited by a wing, has controlled slotted air-feed shutter louvers, a boundary layer drain slot is made between the air intake element and the wing, it contains an integrated complex of avionics, which includes: an information exchange system, an on-board digital flight control and training computer system military operations, external storage device and information input system, inertial system, radio navigation system stations and landings and a transponder of an air traffic control and state recognition system, an automatic radio compass, a radio altimeter with a transceiver and an antenna device, a marker receiver, an integrated aircraft control system with system consoles installed in the pilot and operator’s cabin, armament control system with remote controls installed in the pilot’s and operator’s cabin systems, a comprehensive system of electronic indication, control and aiming, warning information boards installed in crew cabs satellite communication system, two-redundant control system for general aircraft equipment, on-board objective monitoring system, communications radio station, aircraft intercom module, power supply system, external and internal lighting equipment, integrated emergency exit system, two-redundant electronic power plant control system, integrated aircraft control system, which includes: four-time redundant digital computing electro-hydraulic steering gears, four of which are installed in the roll channel and one - in the direction channel, four-fold redundant in the electrical part and twice in the hydraulic control unit, with power and control units, sensors and control panels, four times redundant for the electric part and two times for the hydraulic part parts of electro-hydraulic steering gear of the stabilizer, rotary electro-hydraulic gear of flaps, three-position hydraulic cylinders of gear of wing socks , electro-hydraulic control valves for deflectable wing socks, brake flap drive, mechanically interconnected control rooms of the first and second cabs, trimmer mechanisms along pitch, roll and direction channels, control panels, flexible flap control shafts, planetary gearboxes-release joints flaps, a control system for the brake flap, a control system for general aircraft equipment, which includes: a control panel, a light-signal board with push-buttons, type PKN, a central light signal, multiplex information exchange channel, transformations and calculations unit, control and monitoring unit, airborne equipment systems, air conditioning system, integrated control system, multifunction indicators, inter-aircraft intercom, pneumohydrosystem, wheel braking system, fire protection system, air system lantern visor glass protection, a set of emergency escape facilities, a system for launching, controlling and monitoring a power plant, systems and power supplies, a fuel control and monitoring system, an auxiliary power unit start-up and control system, an oxygen system, a chassis control system, a radio altimeter, a radio compass, a lamp lock alarm system, an external lighting equipment system, a fuel system, a shutter case control system, a centralized oil filling system, guidance and landing equipment, rangefinder, on-board digital computer, integrated aircraft power supply system, which includes : alternating current subsystem consisting of: left drive generator, right drive generator, auxiliary power unit generator, airfield power source connection unit, left and right side alternating current bus, contactors, control, protection and control units, phase rotation unit, blocks current sensors, left rectifier device, right rectifier device, DC subsystem consisting of: left battery, right battery, emergency emergency bus left о and starboard side, disconnectable DC buses of the left and starboard side, contactors, protection and control devices of rectifier devices, electric starter of auxiliary power unit, electronic display complex, which includes: the first control board digital computer, multiplex communication channel, local circular channel of information exchange, left, central and right multifunctional digital indicators, kallimotor aviation indicator, multifunctional control ults, power supply system with the generator of the main power plant, generator of the auxiliary power plant, battery, sensor system of operational parameters, the second control on-board digital computer, inter-machine communication channels, the first (left) and second (right) local radial channels of information exchange backup, television signal switching unit, universal device for mounting external suspended objects on an airplane, which includes : a pylon with wing mounts, lockless mounts, a holder lock, compartments for accommodating disconnect and / or launch units, a hatch, a removable tail fairing, an autonomous on-board oxygen system, which includes: an overload suit, an oxygen system of the ejection seat, a sensor breathing air supply, non-return valve, oxygen hose, pressure switch, pressure regulator, oxygen hoses, oxygen device, oxygen mask, safety helmet, cabin valve, surge hose lines, on-board oxygen production unit, hoses, air preparation device. 2. The aircraft according to claim 1, characterized in that the longitudinal axis of the auxiliary power unit is directed along the axis of the aircraft, and the exhaust pipe is brought out under the fuselage to the side. 3. The aircraft according to claim 1 or 2, characterized in that the channel of each air intake in the front part is inclined upward, and at the engine level goes into a horizontal position.

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