Composite materials in the aircraft industry. What is an office chair made of What material is used in helicopter seats

Cushions for chairs and sofas.

Aviation seat cushions are made from a soft material called polyurethane foam or foam rubber. Easier - PPU.

Aviation seat cushion foam rubber is a soft aviation non-combustible material (tested by special tests for fire safety) intended for use in a passenger aircraft cabin, which does not have vents and windows designed to ventilate the room in the event of a pillow fire.

In accordance with aviation regulations, a foam rubber pillow, dressed in a decorative (and possibly also an additional protective) cover made of non-combustible fabric, is subjected to fire tests for the second time together with covers in a special laboratory to determine the combustibility of the product assembly.

In the cabin of a passenger aircraft, only those pillows that meet the requirements of aviation regulations should be used, which is confirmed by the test report and the quality stamp of a certified aviation pillow manufacturer.

In the case of application household foam rubber for the manufacture of aircraft seat cushions, testingthis pillow will not pass, the fire in the aircraft spreads instantly, and when burning household foam rubber, toxic products are released (xylene, Toluene diisocyanate ), the number of which exceeds the permissible norms from 3 to 65 times, which can lead passengers and crew members to diseases of varying severity.

Unfortunately, sometimes there are cases when airlines use pillows made of household foam rubber microporki for shoes, rubber – combustible and hazardous materials. Even in protective covers made of non-combustible fabric, these pillows will burn out instantly. In this case, the chances of a passenger surviving a fire are negligible.

FORBIDDEN!

In these cases, documents confirming the airworthinesspillows and permission to install them on a seat, the airlines do not have.

However, pillows don't last forever. During prolonged use, the pillow loses its shape and becomes flat, the foam rubber breaks and falls apart.

Every time a passenger sits on a torn pillow, a stream of small, invisible to the eye particles of foam rubber enters the air environment of the passengersalon. And passengers, both adults and children, breathe this air without even knowing it.

To breathe or not to breathe?

To improve the flight and tactical characteristics of combat aircraft and helicopters in the countries of the aggressive bloc, expensive programs are being carried out to reduce the weight of the aircraft structure through the use of new, more promising materials, which include the so-called composite materials.

The leading place in the capitalist world in the development of composite materials and their use in the construction of aircraft (especially for military purposes) belongs to, where the pace of work in this area is constantly growing. If in 1958 the Pentagon was allocated $400,000 for R&D to create such materials, then by 1967, about $11 billion was spent on the same item. The coordination of ongoing research (in relation to aircraft structures) is carried out by the US Air Force Materials Laboratory and. The Materials Laboratory is engaged in evaluating the effectiveness of the use of composite materials in the construction of military aircraft. Currently, under contracts with the Air Force and programs financed by large aircraft manufacturers, a large number of structural elements of aircraft and helicopters from composite materials are being produced and tested.

A composite material (sometimes called a composite) consists of a high-strength filler oriented in a certain direction and a matrix. Fibers of beryllium, glass, graphite, steel, silicon carbide, boron or the so-called whiskers of aluminum oxide, boron carbide, graphite, iron, etc. are used as reinforcing fillers (the power base of the composition). Matrices are made of synthetic resins (epoxy, polyester, organosilicon) or metal alloys (aluminum, titanium and others) The connection of fibers or whiskers with a matrix is ​​carried out by hot pressing, casting, plasma spraying and some other methods.

Composite materials based on high-strength fibers are most widely used in aircraft and rocket building abroad. A composite material behaves as a single structural whole and has properties that its constituent components do not have. A feature of composite materials is the anisotropy of their properties (that is, the dependence of the physical, including mechanical, properties of materials on the direction), which is determined by the orientation of the reinforcing fibers. The given strength of the material is obtained by orienting the filler fibers in the direction of the main force. Foreign experts believe that this opens up new possibilities in the design of power elements of aircraft and helicopters.

According to foreign experts, from the point of view of the characteristics of specific strength and specific rigidity, the most promising are composite materials in which boron, boron carbide and carbon fibers are used as reinforcing reinforcement. These materials include boron epoxy materials (boroplastics, carbon fiber reinforced plastics, boron aluminum).

Boron epoxy composite materials

Abroad, the most widely used materials (boroplastics) with a reinforcing filler of boron fibers (boron fibers) and epoxy matrices. According to the foreign press, the use of boron plastics makes it possible to reduce the weight of the structure by 20-40%, increase its rigidity and increase the operational reliability of the product. Composite materials based on boron fiber have high strength, stiffness and fatigue resistance. For example, in the foreign press it was noted that the ratio of the specific strength of boron plastics to the specific strength of an aluminum alloy in tension is 1.3-1.9, compression - 1.5, shear - 1.2, crushing - 2.2, and the fatigue characteristic increases 3.8 times. In addition, boroplastics retain their qualities in the temperature range from -60 to + 177°C. The combination of these properties predetermined the prospects for the wide use of boroplastics in aviation and rocket and space technology.

As follows from the report of the foreign press, the scale of the use of boroplastics in the US aircraft industry is already very significant at the present time. For example, one fighter consumes about 750 kg of boron plastics. These materials are used to reinforce the elements of the power set with boron plastic overlays, which reduces the weight of structural elements and increase their load-bearing capacity, as well as for the manufacture of skins.

Thanks to the use of boroplastics, the production technology is greatly simplified, and, in addition, it is possible to reduce total components and parts in some elements of the aircraft structure. For example, according to McDonnell Douglas specialists, in the manufacture of the F-4 aircraft rudder from boroplastics, the number of parts was reduced from 240 to 84.

Composite materials with carbon fibers

Foreign experts believe that under conditions of high temperatures that occur during supersonic flight, composite materials based on matrices reinforced with graphite (carbon) fibers are the most effective. The use of these materials in the construction of modern and advanced supersonic aircraft is beneficial in terms of saving the weight of the structure, especially for components whose weight is more determined by the requirements of rigidity than strength. Materials with carbon fibers based on epoxy matrices (CFRP) and materials based on graphitized carbon matrices reinforced with carbon fibers ("carbon-carbon") have received the greatest distribution abroad.

CFRP

Foreign press notes that carbon plastics have a low specific gravity - 1.5 g / cc. (aluminum alloys 2.8 g/cc, titanium alloys 4.5 g/cc); high rigidity, vibration strength and fatigue strength. All this makes them one of the most promising materials for the production of aviation and space technology. It is reported that for all main types of acting loads, the specific strength of carbon plastics is higher than the strength of an aluminum alloy. Foreign experts note that the strength and rigidity of carbon fiber is about six times higher than that of the main grades of steel used in aircraft structures.

In 1969, the US Air Force Materials Laboratory awarded Northrop a contract to develop prototype graphite-based composite structures. Initially, the use of carbon fiber in aircraft structures was insignificant due to the high cost of carbon fiber (700-900 dollars per 1 kg). Subsequently, as a result of the organization of a wide production of fiber, the cost decreased to 120-150 dollars. But according to the forecasts of American experts, in three to five years it will not exceed 50-80 dollars.

According to the foreign press, at present the use of carbon fiber in the aircraft industry has increased significantly. Various structural elements made of this material are being tested on F-5E, A-4D and F-111 aircraft. The Boeing company, under a contract with the US Air Force, is exploring the possibility of using these materials in the wing design of a promising high-altitude unmanned reconnaissance aircraft. Similar work is being carried out in other capitalist countries. For example, the British firm "British Aircraft", under a contract concluded with the British Ministry of Defense, creates elements of the airframes of some aircraft from carbon fiber.

Composite materials "carbon-carbon" have a low specific gravity (1.4 g/cm3), high heat-shielding properties, the ability to maintain strength characteristics at temperatures above 2500 degrees Celsius. Due to these and other qualities, they are considered very promising for the manufacture of those parts and assemblies of aircraft that operate at high temperatures, as well as for heat shields of aircraft, primarily spacecraft. According to foreign press reports, wheel brake parts are currently developed from this material for aircraft, their weight is about 30% of the weight of steel brakes. According to specialists from the American company Dunlop, the service life of brake devices made of these materials is 3,000 landings, which is five to six times longer than the service life of conventional brakes.

Boron aluminum composite material (boron aluminum)

Boron fibers (sometimes coated with silicon carbide) are used as a reinforcing filler for this composite material, and aluminum alloys are used as a matrix. Boro-aluminum is 3.5 times lighter than aluminum and 2 times stronger than it, which allows you to get significant weight savings. In addition, at high temperatures(up to 430°C) boron-aluminum composite material has 2 times big values specific strength and stiffness in comparison with titanium, which makes it possible to use it for aircraft with flight speeds of M = 3, in the designs of which titanium is currently used. Foreign experts consider boron-aluminum also one of the promising composite materials, the use of which can save up to 50% of the weight of the aircraft structure.

According to foreign press reports, work on the study of the characteristics of boron aluminum and its introduction into the aircraft industry is being carried out by several American firms. For example, the General Dynamics company manufactures structural elements of the tail section of the F-111 aircraft from this material, and the Lockheed company produces an experimental caisson of the center section of the C-130 aircraft. Boeing specialists are studying the possibility of using boron-aluminum material in stringers of superheavy aircraft.

At present, boron-aluminum composite material is increasingly used in the construction of aircraft engines. According to the foreign press, the Pratt-Whitney company uses it in the production of fan blades for the first and third stages of the JT8-D, TF-30, F-100 turbofan engines, and the General Electric Company uses the fan blades of the J-79 engine, which, according to the company's specialists, it will allow to obtain about 40% savings in the weight of these elements.

There are 79 programs in the United States, within the framework of which work is being carried out on the research and practical use of composite materials in the aircraft industry.

Analyzing the results obtained during the performance of experimental work, foreign experts believe that composites can be used in the design of most components and parts of a combat aircraft. On fig. 1 shows a diagram of the airframe of a combat aircraft indicating those elements in the designs of which, according to the views foreign specialists, the use of composite materials is possible.

Rice. 1. Scheme of a combat aircraft airframe made using composite materials: 1 - cockpit glazing frame; 2 - cabin trim; 3 - main spars; 4 - power set of the wing and tail; 5 - pylon; 6 - fuselage skin; 7 - slats; 8 - flaps, spoilers, ailerons: 9 - rudders and elevators; 10 - engine attachment points and hatches; 11 and 12 - cabin floor structure; 13 - front and rear walls of the cabin; 14 - the main elements of the transverse power set; 15 - beams;: 16 - fuel tank.

Created by Rockwell International strategic bomber B-1 internal and external spars located in the rear fuselage are made using overlays of boron epoxy composite material. These spars consist of solid boron plastic linings connected to metal parts. Metal elements (steel, titanium) provide strength, and boron plastic linings increase the rigidity of the spars. It is noted that the spars of this design not only have improved mechanical properties, but also 28-44% lighter than all-metal ones.

Envisaging the further introduction of composite materials into the design of the B-1 bomber, the US Air Force Materials Laboratory signed contracts with Rockwell International to develop a keel from graphite-epoxy and boron-epoxy materials, and with Grumman to create an aircraft stabilizer from these materials.

In accordance with the program implemented by General Dynamics (under contract with the US Air Force), epoxy boron plastic reinforcing pads are installed on the high-strength steel lower surface of the hinged wing support of the fighter-bomber. American experts believe that the use of these overlays more than doubles the fatigue strength of the hinge joint of the wing turning unit. On two F-111A aircraft, experimental stabilizers made of boron-epoxy composite material are being tested, which, according to foreign press data, are 27% lighter than conventional ones.

In the F-l4 aircraft, the use of composite materials in the load-bearing structure was envisaged at the very beginning of its design. Four stabilizer skin panels are made from a composite material based on boron fiber.

According to the foreign press, the results of the tests showed that the fatigue characteristics of the stabilizer with boron plastic sheathing are 2.5 times higher than the specified technical requirements, but at a cost it is currently equivalent to an all-metal one. The total weight of the stabilizer with boroplastic casing is 350 kg; weight savings compared to a titanium-clad stabilizer of 82 kg (or 10%). Compared to a stabilizer of a similar design made of aluminum alloys, the weight gain is even greater - 117 kg (27%).

In the design of the F-15 aircraft (McDonnell Douglas), based on considerations of ensuring the required centering in order to save the weight of the tail section of the aircraft, the skin of the horizontal controllable stabilizers and vertical tail unit is made of boron plastic. According to foreign press reports, fatigue tests of the F-15 airframe with composite skin panels have been completed. The test duration is 10 thousand hours, which is four times its normal resource. Then, static tests of the horizontal controllable stabilizer were carried out at a load twice the design destructive load; The stabilizer has withstood these tests. Compared to a titanium horizontal stabilizer design, the weight savings with boron skins was 22%.

As noted in the foreign press, the F-15 aircraft is the first US Air Force military aircraft to have a Goodyear brake system, the parts of which are made using a composite material based on carbon fibers. This provided, according to American experts, weight savings (about 32 kg for each brake) and smoother and at the same time more efficient braking, and also increased the reliability of the brake system.

McDonnell Douglas has been conducting research for the third year under a special program that provides for the use of composite materials for various elements of the F-15 wing, which, according to the calculations of the company's specialists, will reduce the weight of the wing by 130-180 kg. In the course of strength tests, an aircraft wing made of composite materials collapsed under a load of 110% of the calculated destructive load. Flight tests of this wing are planned to begin in 1976 (if static tests are successfully completed).

Foreign press reports that the high cost of technical equipment required for the manufacture of parts from such materials did not allow the use of promising composite materials in the proper amount. However, the use of composite materials in the designs of new US combat aircraft is increasing. The experience of using graphite-epoxy composite materials obtained by General Dynamics in the development of the F-111 aircraft was also taken into account when creating the F-16 aircraft. By making the fin skin, stabilizer and rudder out of carbon fiber, the firm has been able to reduce the weight of the F-16's rear fuselage by about 30 percent. Currently, the company, under contract with the Air Force, is developing the front fuselage of this aircraft from graphite-epoxy materials.

During the modernization of the C-5A heavy military transport aircraft, composite materials were used to create some components and parts of the aircraft airframe (for example, slat sections). On fig. 2 shows a slat section made using boron epoxy material and plain metal. The new section has increased strength and rigidity, it is much lighter than the metal one.

Rice. Fig. 2. Section of the slat of a heavy military transport aircraft C-5A: at the top - made using composite materials; bottom - made of aluminum alloys

Attempts are being made to use composite materials in helicopter construction. In particular, in order to study the possibility of manufacturing some of the main structural elements of helicopters from such materials, American and West German firms are conducting a number of development work. According to the foreign press, the American Sikorsky Firm is participating in a program that provides for an increase in fatigue life and an improvement in the dynamic characteristics of the CH-54V helicopter due to the hardening of its tail boom with composite materials. It is reported that as a result of strengthening the stringers with boron epoxy material, the resource of the helicopter airframe increased several times, and the weight decreased by 30% (Fig. 3).

Rice. 3. The use of boron plastic to reinforce the tail boom stringers on a CH-54B heavy helicopter.

It was reported in the foreign press that the US Department of Defense signed a $1.2 million contract with Hughes for the development of blades from composite materials. rotor for a helicopter. According to the company's specialists, the use of composite materials in the design of the blade will reduce its weight, maintain strength characteristics, and achieve relative invulnerability of the blade from bullets. In addition, such blades will have a long resource and low durability, and their production can be arranged on an automated line.

The widespread use of composite materials in the design of the main rotor is also planned as part of the promising HLH program, which provides for the creation of a heavy assault helicopter with a maximum payload of about 30 tons. performance of work under the HLH program, manufactured rotors with rotors, composite materials were used in their design.

On the basis of research conducted by the largest American helicopter company Sikorsky in relation to the CH-53D helicopter, it was concluded that the widespread introduction of composite materials in helicopter structures will become expedient in the 80s. The company's specialists believe that maximum efficiency is achieved when composite materials are included in the design of the helicopter fuselage; at the same time, carbon-based material should be used in the most loaded fuselage elements. The analysis showed that due to the use of composite materials, the weight of the structure of the CH-53D helicopter can be reduced by 18.5%.

Studying the experience of using composite materials in aircraft structures, American experts consider these materials to be very promising for rocket and space technology in terms of weight and mechanical characteristics. According to foreign press reports, in the United States, in the manufacture of missile warheads, it is planned to use composite materials with a carbon fiber matrix, which have high radio transparency. Thermal testing of the nozzle is also reported. rocket engine made entirely of composite materials.

A number of parts of artificial Earth satellites, such as antenna frames, are already being made from carbon fiber in combination with an aluminum honeycomb structure. This provided not only weight savings compared to aluminum construction, but also dimensional stability of the panels, since carbon fiber has an extremely low coefficient of thermal expansion (50 times less than that of metals).

Composite materials are planned to be widely used for the manufacture of some elements of the orbital stage being developed in the United States of the Shuttle transport and space system. In particular, carbon-carbon composite material will be used for thermal protection of the fuselage nose, the lower surface of the fuselage nose, and the leading edge of the wing. Boeing has developed a frame for a liquid jet engine basic propulsion system orbital stage, located in the rear fuselage. It is made of boron epoxy composite material combined with titanium alloy elements. This design, according to the company, will allow, compared with conventional titanium, to achieve weight savings of about 30%.

Studies carried out by a number of American aircraft manufacturing companies under the guidance of the US Air Force Materials Laboratory have shown that the use of composite materials in the construction of military aircraft and helicopters of the 80s will not only significantly reduce their weight and cost, but also increase survivability.

According to the forecasts of foreign experts, by the beginning of the 80s the share of composite materials in the airframe of an aircraft will increase to 50%. This should provide 20-30% weight savings equally for both subsonic and supersonic aircraft. The reduction in the weight of the structure achieved in this case will allow increasing the fuel supply or combat load or reducing the size of the aircraft. Moreover, it is believed that the high strength characteristics of these materials can lead to improved aerodynamic characteristics(by reducing the relative thickness of the profile and lengthening the wing), and ultimately - to improve flight characteristics aircraft.

Armchairs are designed to be placed in them and perform functional duties pilot, passenger accommodation, ensuring a comfortable flight, as well as the tolerance of overloads by the pilot and passengers of the helicopter in the event of an emergency landing.

Our seats are so compact that they fit in almost all cabins.

Armchairs not only meet safety requirements, but also have improved ergonomic characteristics.

When creating the chair, the following goals were achieved:

• weight loss
• cost reduction
• compactness
• maximum ergonomics and comfort
• original design

The armchair has an exclusive, modern design. During the development, new original engineering solutions were introduced. The production process is based on the use of advanced, innovative materials.

The chair is a serial product, has interchangeable components and parts. Seat equipment is easily installed on board the helicopter and is located both in flight and against the flight. Each seat is reliable in operation and, under normal operating conditions, requires minimal operating costs.

The design of the chair withstands high shock loads, with less weight, compared to competitor chairs.

Lightweight chairs provide energy savings, and along with safety - economical operation and high ergonomic characteristics.

The multi-stage safety system of our helicopter seat reduces the possibility of injury to the passenger and contributes to saving his life. The energy absorption technology has a high level of reliability, effectively absorbing impact energy in a severe accident or emergency landing.

energy absorbing helicopter seat, designed for overload up to 30g.

Single action energy absorption element.

In one of the seat modifications, it is possible to install, adjust the degree of absorption of impact energy, depending on the weight characteristics of the passenger (option).

The holding and fixation system consists of: two waist belts, two shoulder belts with inertial coils, a belt fixing lock, a system for adjusting the belts along the length and attachment points for the seat belts.

The chair cushions are designed with minimal displacement (recession) and dynamic feedback of the seated person. Pillows are made of self-extinguishing material in accordance with AP27.853.

The design of the chair provides for the installation of armrests (option).

Implementation high degree chair safety did not affect the main parameters such as low weight, comfort, accessibility and maintainability.

SPECIFICATION

ARMCHAIR CONSISTS OF:

• chair frame
• soft pillows
• Cushioning systems with attachment points
• Shock absorber adjustment system depending on the weight of the passenger (option)
• armrests (option)
• Headrest
• Tethered system
• Power supply (option)
• literary pocket
• Case (textile/leather) with pre-selected color scheme

SERVICE

Quick release elements:

• softness
• Cases

Knots with adjustment applied:

• Armrest

HELICOPTER Glider and Cockpit Equipment

1. GENERAL

The fuselage is an all-metal semi-monocoque of variable section, consisting of a frame and skin. The fuselage is the base to which all the units of the helicopter are attached, it houses the equipment, crew and payload.

The design of the fuselage provides its operational dismemberment, which simplifies the repair and transportation of the helicopter. It has two constructive connectors (see Fig. 2.16) and includes a nose and a central part, a tail boom and an end boom with a fairing.

The main construction materials are: D16AT clad duralumin sheet of 0.8 mm thick sheets of which the outer skin is made, hardened B95 duralumin and magnesium alloys.

In the design of many units, stampings from aluminum alloys, castings from steel and non-ferrous alloys, as well as extruded profiles are used. Individual components and parts are made of alloyed steels.

Synthetic materials are used for soundproofing and finishing of cabins.

2. FORWARD FUSELAGE

The forward part of the fuselage (Fig. 2.1), which is the cockpit, is a compartment 2.15 m long, which contains the pilot's seats, helicopter and engine controls, instrumentation and other equipment. Its front part forms a lantern that provides visibility to the crew. The crew cabin is separated from the cargo compartment by frame No. 5H with a door.

Sliding blisters 2 are located on the right and left. In the cabin ceiling there is a hatch for access to the power plant, which is closed with a lid that opens upwards. On the floor of the cockpit there are helicopter control levers and pilots' seats, and a flight engineer's seat is installed in the entrance door to the cockpit. Behind the seats between frames No. 4N and 5N there are battery compartments and shelves for radio and electrical equipment.

The bow frame consists of five frames No. 1N - 5N, longitudinal beams, stringers, stamped stiffeners and a canopy frame. Technologically, the bow is divided into the floor, side panels, ceiling, canopy, sliding blisters and frame No. 5H.

The floor of the cockpit (Fig. 2.2) of riveted construction consists of a set of lower parts of frames, longitudinal beams and stringers. The power frame is fastened with corner profiles and reinforced with profiles and diaphragms in the places of cutouts and fastening of the units.

The flooring and outer skin made of duralumin sheets are attached to the frame. On top of the flooring along the axis of symmetry, between stringers No. 3, two sheets of corrugated duralumin are installed.

In the floor and the outer skin of the floor, hatches were made for mounting the units, access to the nodes and joints of the helicopter control system rods, to the attachment points of the front landing gear, docking bolts of frame No. 5H and pipes of the heating and ventilation system.

In the outer skin between frames No. 2N and ZN, hatches 10 were made for the installation of MPRF-1A landing and taxiing lights. On Mi-8P helicopters, under the floor of the cockpit between frames No. 4N and 5N, a second flashing beacon MSL-3 is installed.

Rice. 2.2. Cabin floor forward fuselage:

1, 5, 6, 11 - openings for helicopter controls; 2 - hole for electrical wiring of the dashboard; 3 - overlays; 4 - hole for the pipe of the heating system; 7 - hatch for approaching the shock absorber of the front landing gear; 8 - assembly and inspection hatches; 9 - a hatch for a flashing beacon; 10 - hatches for headlights.

To protect the flooring from wear, four pads 3 made of delta wood are installed under the directional control pedals. Brackets for attaching seats, helicopter control units, instrument panels and an autopilot console are mounted on the floor.

The side panels are made of stamped stiffeners, profiles and duralumin sheathing. Stamped stiffeners, together with cast magnesium profiles, form the frames of the openings for the right and left sliding blisters.

Rubber profiles are installed along the front and rear edges of the openings to seal the cockpit. Outside, above the openings and in front of them, gutters for water drainage are attached. In the upper part of the frame sealing of the openings, mechanisms for emergency blisters ejection are mounted from the inside.

On the right and left sides between the frames No. 4H and 5H, compartments are made to accommodate batteries (two on each side). The compartments are closed from the outside with covers that are locked with screw locks. The lids are hinged and, for ease of use, are held in a horizontal position by two steel rods. Guides are installed in the compartments along which containers with batteries move. The internal surfaces of the battery compartments are pasted over with heat-insulating material. Under the blisters between frames No. 1H and 2H, BANO-45 navigation lights are installed. On the left side, in front of the battery compartments, there are cutouts for airfield power plug connectors 4 (see Fig. 2.1).

The ceiling of the cockpit is made of stamped stiffeners, a longitudinal and transverse set of diaphragms, profiles and duralumin sheathing. The skin is riveted to the frame with special spiked rivets to prevent legs from slipping when servicing the power plant.

There is a hatch in the ceiling for access to the power plant. The design of the hatch and cover provides protection against water ingress into the cockpit.

The riveted manhole cover is mounted on two hinges 1 (Fig. 2.3). A spring-loaded latch is mounted in the first hinge, which automatically locks the lid in open position. When the cover is opened, the profiled rib 10 with its beveled section depresses the axis of the latch 13 until the axis, under the action of the spring 12, passes to the straight section of the rib, after which the hatch cover is locked.

Rice. 2.3. Access hatch to the power plant:

1 - hatch hinges; 2 - stops; 3 - latch button; 4 - fork; 5 - adjusting clutch; 6 - shaft, 7 - latch; 8 - hook; 9 - handle; 10 - profiled rib; 11 - locking pin; 12 - spring; 13 - latch.

When closing the manhole cover, you must first press the protruding end of the latch and move the axle beyond the profiled edge of the hinge loop. In the closed position, the hatch cover is fixed with a lock. The lock mechanism consists of a handle 9 with a locking device, a fork 4, an adjusting clutch 5 and a shaft with two legs 6. When opening the hatch cover, press the latch button 13, disengage the latter from the hook 5, then turn the handle down. In this case, the shaft will turn clockwise, and the paws will release the cover. There are two viewing windows in the hatch cover for visual observation in flight of the state of the engine air intake inlet tunnels. The sealing of the hatch in the closed position is provided by rubber gaskets, which are pressed by a special profile attached to the hatch around the perimeter. In case of violation of the tightness of the hatch, the elimination is carried out by the adjusting clutch 5 of the lock control rod.

Frame number 5H. The forward part of the fuselage ends with a docking frame No. 5H (Fig. 2.4). The frame is a duralumin wall edged along the perimeter with a pressed corner profile, the end beam of which forms a flange for joining with the central part of the fuselage. The wall is reinforced with a longitudinal and transverse set of angle profiles. Along the axis of symmetry in the wall of the frame, an opening was made for the front door to the cockpit. The opening is edged with a pressed duralumin corner, to which a rubber profile is fixed with screws.

Shelves for equipment installation are attached to the front wall of the frame on both sides of the doorway. In the left part of the wall at the top and bottom there are holes for the passage of rods and cables for controlling the helicopter. On the right and left sides of the wall of frame No. 5H, special plates are installed on the side of the cargo compartment to ensure flight safety. A casing with removable covers is attached to the rear left side of the wall of frame No. 5H, enclosing the system of rods and rockers for controlling the helicopter and electrical equipment harnesses. A folding seat is attached to the casing. In the transport version, on the right side of the doorway on the side of the cargo compartment, a box is riveted to the wall, in which containers with batteries 3 are placed (see Fig. 2.1). The box is equipped with guides and is closed with lids with screw locks.

The cockpit door is made in the form of a duralumin plate. It is hung on hinges and equipped with a lock with two handles, and two locks - valves - are installed on the side of the cockpit. An optical micro-peephole is installed at the top of the door. In the doorway between frames No. 4H and 5H, a folding seat of the on-board technician with seat belts is installed.

The cockpit canopy consists of a frame and glazing. The frame of the lantern is assembled from duralumin profiles, stiffeners and facing frames, fastened together with screws and rivets.

Rice. 2.4. Frame No. 5H

The lantern is glazed with oriented organic glass, with the exception of two front windshields 1 (see Fig. 2.1) (left and right), made of silicate glass, which are electrically heated and equipped with wipers. Along the perimeter, the glass is edged with rubber profiles, inserted into cast magnesium frames and pressed through the duralumin cladding with screws with special nuts. After installation, for tightness, the edges of the frames inside and outside are coated with VITEF-1 sealant.

The blister (Fig. 2.5) is a frame cast from magnesium alloy, into which a convex organic glass 14 is inserted. The glass is fixed to the frame with screws through the duralumin lining 11 and a rubber sealing gasket. The blisters are equipped with handles 12 with lockable pins 7 connected to the levers 13 by cables 8. The left and right blisters can only be opened from the cockpit.

The blisters are moved back along the upper and lower guides made of special profiles.

Upper internal guide profiles 5 are mounted on balls which are located in steel cages. The outer U - shaped guide profile 6 has brackets with lugs for the locking pins of the blister emergency release mechanism and drilling with a pitch of 100 mm for pin 7 of the lock to fix the blister in extreme and intermediate positions. In the lower part of the blister frame there are grooves in which the lower guide profiles 9 slide along the felt pads, fixed with screws to the opening frame.

Each blister can be dropped in an emergency using the handle located above the blister inside the cockpit. To do this, the handle must be pulled down, then under the action of springs 1, the locking pins 2 will come out of the lugs of the brackets 3, after which the blister must be pushed out. In the lower profiles of the frames of the openings there are slots for supplying hot air to the blisters. On the left blister, a visual icing sensor is installed at the bottom.

Rice. 2.5. Sliding blister:

1 - spring; 2 - locking pin; 3 - bracket; 4 - handle for emergency release of blisters; 5 - internal guide profiles; 6 - outer guide profile; 7 - pin; 8 - cable; 9 - lower guide profiles; 10 - felt pad; 11 - lining; 12 - handle; 13 - lever; 14 - glass; 15 - outer handle of the blister.

3. CENTRAL FUSELAGE

General information. The central part of the fuselage (Fig. 2.6) is a compartment located between frames No. 1 and 23. It consists of a frame, working duralumin skin and power units. The frame consists of a transverse and longitudinal set: the transverse set includes 23 frames, including frames No. 1 and 23 - docking, frames No. 3a, 7, 10 and 13 - power, and all other frames of lightweight construction (normal). The longitudinal set includes stringers and beams.

Frames provide a given shape of the fuselage in cross section and perceive loads from aerodynamic forces, and power frames, in addition to the above loads, perceive concentrated loads from helicopter units attached to them (chassis, power plant of the main gearbox).

Technologically, the central part is assembled from separate panels: cargo floor 15, side panels 3.5 and ceiling panel 4, rear compartment 7.

Rice. 2.6. The central part of the fuselage:

1 - attachment point of the shock absorber of the front landing gear; 2 - sliding door; 3 - left side panel; 4 - ceiling panel; 5 - right side panel; 6 - attachment point of the shock absorber of the main landing gear; 7 - rear compartment; 8 - cargo hatch doors; 9 - attachment point of the strut of the main leg of the chassis; 10 - mount of the axle shaft of the main leg of the chassis; 11, 12, 13, 14 - attachment points of the external fuel tank; 15 - cargo compartment floor panel; 16 - attachment point of the strut of the front leg of the chassis.

a - a hole for the air intake pipe from the cargo compartment; b - hole for the pipeline of thermal air; c - hole for the duct of the heating and ventilation system; g - spare nodes; d - attachment points for the tie-down bands of outboard fuel tanks; e - attachment point of the mooring device.

In the central part, between frames No. 1 and 13, there is a cargo cabin, ending at the rear with a cargo hatch, and between frames No. 13 and 21 there is a rear compartment with cargo flaps 5. Behind frame No. 10 there is a superstructure that smoothly passes into the tail boom. In the passenger version, the compartment between frames No. 1 and 16 is occupied by the passenger compartment, behind which the luggage room is located. Engines are located above the cargo compartment between frames No. 1 and y, and the main gearbox is located between frames No. 7 and 10. In the superstructure between frames No. 10 and 13 there is a consumable fuel tank, and between frames No. 16 and 21 - a radio compartment.

Rice. 2.7. Frames of the central part of the fuselage:

a - power frame No. 7; b - power frame No. 10; c - power frame No. 13; g - normal frame; 1 - upper beam; 2 - side part; 3 - fitting; 4 - lower part; 5 - arched part; 6 - mooring ring.

All other frames, except for docking frames, are made composite, including the upper part, two side and lower parts. These parts of the frames, as well as the stringers, are included in the design of the panels and, during assembly, the parts of the frames are joined together, forming a load-bearing frame of the central part of the fuselage.

The most loaded elements of the central part of the fuselage are power frames No. 7, 10 and 13, as well as the floor panel. Power frames No. 7 and 10 (Fig. 2.7) are made of large forgings of the AK-6 alloy, pressed and sheet parts, which form a closed profile, including the upper beam 1, two sidewalls 2 and the lower part 4.

The upper beam consists of two parts connected by steel bolts in the plane of symmetry. At the corners of the beams there are holes for the bolts of the frame of the main gearbox.

Docking of the upper beam of frame No. 7 with the sidewalls was carried out using milled combs and two horizontally located bolts, and the docking of the sidewalls of frame No. 10 with the upper beam was made using a flange and vertically located bolts. The lower parts of frames No. 7 and 10 consist of walls and 4 corners riveted to it, forming an I-profile in cross section. At the ends of the beams, docking fittings 3 stamped from AK-6 alloy are installed, with which the lower beams of the frames are joined to the sidewalls with steel bolts.

On the outer part of frame No. 7, steel attachment points for external fuel tanks are installed on both sides. On frame No. 10, combined units are installed for simultaneous fastening of the suspension struts of the main landing gear and mooring devices. In addition, in the lower part of the frame on both sides there are rear attachment points for outboard fuel tanks.

Frame No. 13 of riveted design is made of sheet duralumin and extruded corner profiles. The lower part of the frame is made of three forgings of AK-6 alloy, bolted together. With the sidewalls of the frame, the lower part is riveted with the help of fittings, in which there are holes for installing mooring rings 6. An inclined frame is attached to the bottom of the frame No. 13, which closes the cargo compartment and is the power edging of the cargo hatch. It has two nodes on each side for hanging cargo flaps.

In the upper part of frame No. 13, an arched part 5 is installed, which is part of the fuselage superstructure; it is stamped from sheet duralumin and has notches for the passage of stringers.

Lightweight (normal) frames (see Fig. 2.7) are similar in design and have a Z-shaped profile in cross section. The upper and side parts of the frames are stamped from sheet duralumin and are butted together with overlays. The frames are reinforced with an angle profile along the inner contour, and notches for stringers are made along the outer contour.

The lower parts of normal frames have upper and lower chords made of angle and tee profiles, to which a wall of sheet duralumin is riveted. Fittings stamped from AK-6 alloy are riveted at the ends of the lower parts of the frames, with the help of which they are riveted to the sidewalls of the frames.

Outside, on the starboard side on frame No. 8, on the left side between frames No. 8 and 9, as well as on frame No. 11, and on both sides there are dee nodes for attaching tapes of outboard fuel tanks. From below, along the lower parts of the frames, overhead nodes made of ZOHGSA steel are installed for attaching the chassis. On the frame No. 1, along the longitudinal axis of the helicopter, the attachment point of the front suspension strut is installed, and on the sides of the frame and the longitudinal beams of the floor, nodes with spherical nests are riveted under the jack supports. On frame No. 2, attachment points for the front landing gear struts are installed. On frame No. 11, attachment points for axle shafts are installed, and on frame No. 13, attachment points for struts of the main landing gear are installed.

In the ceiling panel between frames No. 7 and 13, as well as in the side panels, there are stringers made of special D16T duralumin corner profiles with chamfers to improve gluing with the skin. The remaining stringers are installed from corner profiles.

The cargo floor (Fig. 2.8) of a riveted structure consists of the lower parts of the frames, longitudinal beams 11, stringers, flooring made of corrugated sheet 338 AN-1 and outer duralumin sheathing. The middle longitudinal part of the flooring, located between frames No. 3 and 13, is reinforced with transverse rigid elements and fastened with screws with anchor nuts to special longitudinal profiles. Corner profiles made of duralumin sheet D16AT and L2.5 are riveted on top of the flooring along the sides of the floor, with the help of which the side panels are connected to the floor of the cargo compartment. Floor loading zones from transported wheeled vehicles are reinforced with two longitudinal trough-shaped profiles. To secure the transported cargo on the floor along the sides, 27 mooring knots 5 are installed.

The frames and beams in the places where the mooring units are installed have stamped brackets and fittings made of AK6 alloy. On frame No. 1 along the axis of symmetry of the cargo floor there is a node 1 for fastening the rollers of the LPG-2 electric winch when pulling loads into the cabin. At the installation site of the LPG-2 electric winch on the wall of the longitudinal beam

a stamped fitting made of AK6 alloy is reinforced, in the flange of which there are two threaded holes for the plate 2 fastening bolts under the base of the LPG-2 electric winch. On the floor between frames No. 1 and 2, a casing is installed to protect the rollers and cables of the LPG-2 electric winch, and in the opening of the sliding door there are two holes for fixing a removable entrance ladder.

In the walls of the longitudinal beams of the cargo floor at frame No. 5, as well as in the wall of frame No. 1 at the starboard side, there are holes for pipelines 12 of the heating and ventilation system of the cabins. The walls around the holes are reinforced with stamped edgings made of AK-6 alloy. On the left and right sides of the floor between frames No. 5 and 10 there are cradles for additional fuel tanks.

Rice. 2.8. Cargo cabin floor panel:

1 - attachment point for electric winch rollers; 2 - plate under the base of the electric winch; 3 - mooring knots; 4 - hatch for the ARC-9 antenna; 5, 8 - hatches to the shut-off valves of the fuel system; 6 - mounting hatch; 7 - hatch to the latch of the cable for cleaning the external suspension; 9, 17, 23 - technological hatches; 10 - hatch for the ARK-UD antenna; 11 - floor frame beams; 12 - pipeline of the heating system; 13 - attachment points of the struts of the shock absorber of the front landing gear; 14 - a niche for the frame of the ARK-9 antenna; 15 - cutouts for pipelines of additional fuel tanks; 17 - attachment points of the external suspension; 18 - supports for hydraulic lifts; 19 - attachment points of the struts of the main landing gear; 20 - hatch control connections pipelines of the fuel system; 21 - attachment points of the semi-axes of the main landing gear; 22 - attachment point of the shock absorber of the front landing gear.

In the cargo floor between frames No. 5 and 6, attachment points for the ARK-9 loop antenna are installed, and between frames No. 8 and 9, attachment points for the antenna amplifier and the ARK-UD antenna unit are installed.

There are assembly and technological hatches in the flooring, closed with covers on screws with anchor nuts. Along the axis of symmetry in the removable part of the flooring there are hatches 4 for inspection and access to the ARK-9 loop antenna, fuel valves 5 and 8, the ARK-UD antenna unit and antenna amplifier and the handle for fixing the external suspension in the retracted position.

On Mi-8T helicopters of the latest series, in the cargo floor between frames No. 8 and 9, a hatch was made for the passage of external cable suspension lines with a load capacity of 3000 kg.

When working with an external suspension, the hatch has a fence. Cable external suspension nodes are located inside the cargo compartment on the upper beams of frames No. 7 and 10. In the stowed position, the suspension rises to the ceiling of the cargo compartment and is fastened with a DG-64M lock and a cable to a special bracket installed between frames No. 10 and 11. Cargo slings fit into cargo box. The guard is folded and, with the help of rubber shock absorbers, is attached behind the back of the landing seat in the left cargo flap. The hatch in the floor of the cargo compartment is closed by paired (internal and external) covers from the cargo compartment.

The side panels (see Fig. 2.6) are riveted from the side parts of (normal) frames, stringers from corner profiles and duralumin sheathing. The rear parts of the panels end with an inclined frame. On the right and left panels there are five round windows with convex organic glass, except for the first left window glazed with flat organic glass. The glasses are fixed to the cast magnesium frames with screws with special nuts and sealed along the contour with rubber gaskets, and the edges of the frames are coated with sealant inside and out after the glass is installed.

On the left side of the panel between frames No. 1 and 3 there is an opening for sliding door 2, edged with a frame of duralumin profiles. On the upper part of the doorway on the side of the cargo compartment, knots for a rope ladder are installed, and a gutter for water drainage is attached above the doorway.

The door (Fig. 2.9) of a riveted structure is made of a frame and outer and inner skins riveted to it, installed on the lower and upper guides, along which it slides back on balls and rollers. The upper guide 11 is a U - shaped profile, in which the skid 14 and two rows of balls 12 are installed. Brackets 15 are riveted to the skid, which are connected to the door by locking pins 13 mounted on the door. In the open position, the door is held by a spring latch mounted on the side of the fuselage from the outside.

Rice. 2.9. Sliding door:

1 - latch; 2 - pin spring; 3, 4 - handles for emergency reset of the door; 5 - cable; 6 - glass; 7 - inner door handle; 8 - springs; 9 - heck; 10 - outer door handle; 11 - top guide; 12 - ball bearings; 13 - locking pin; 14 - skid; 15 - bracket; 16 - roller.

The door has a round window with flat organic glass and is equipped with two locks. A key lock with two handles 10 and 7 (external and internal) is installed on the front edge of the middle part of the door.

A pin lock is mounted in the upper part of the door, for emergency dropping of the door, with inner and outer handles 3 and 4. The upper lock is connected with the middle lock by cable wiring and when the upper lock is opened, the middle lock is opened simultaneously. In case of emergency dropping of the door, it is necessary to turn the outer or inner handle back in the direction of the arrow, while the locking pins 13 of the upper lock come out of the holes of the brackets, and the latch 9 of the middle lock is disengaged by cable 5, after which the door should be pushed out.

To prevent spontaneous opening of the door in flight, a device is installed on it that locks the door in the closed position.

The ceiling panel (Fig. 2.10) consists of the upper parts of the frames, stringers and sheathing, riveted together. In lightweight (normal) frames, notches were made for the passage of stringers, and along frames No. 3, 3a, 7, 10, the stringers were cut and joined through toothed belts made of duralumin sheet. The lining of the ceiling panel between frames No. 1 and 10 is made of sheet titanium, and between frames No. 10 and 13 is made of duralumin sheet. In the lining of the ceiling panel between frames No. 9 and 10, holes are made for the angles of the fire hydrants of the fuel system, and between frames No. 11 and 12 - hatch 6 for the fuel pumps of the supply tank. Gutters made of extruded profiles are installed on the casing and holes are made for drainage pipelines for water flow.

On top of the frames of the ceiling panel, nodes are installed: on frame No. 3 - four nodes 1 for mounting engines, on frames No. 5 and 6 - nodes 2 and 3 for fastening the engine fixation device with the gearbox removed, on frames No. 6 and 7 - nodes 5 for fastening frame No. 1 of the hood, knot 4 of fastening of the struts of the hood and fan.

The rear compartment 7 (see Fig. 2.6) is a continuation of the central part of the fuselage and, together with the cargo flaps, forms the rear contours of the fuselage. The rear compartment of the riveted structure consists of the upper arched parts of the frames, stringers and outer skin.

Technologically, the compartment is assembled from separate panels and is a superstructure located on top of the cargo compartment, smoothly turning into the tail boom. The superstructure ends with a docking frame No. 23.

At the top between frames No. 10 and 13 there is a container for a consumable fuel tank. Between frames No. 16 and 21 there is a radio compartment, in its lower part between frames No. 16 and 18 a hatch is made for entering from the cargo compartment into the radio compartment and into the tail boom.

On frames No. 12, 16 and 20, fittings are installed at the top for the supports of the transmission tail shaft. Docking of the rear compartment with the ceiling and side panels is carried out with corner profiles and external linings.

The skin of the central part of the fuselage (Fig. 2.11) is made of D16AT duralumin sheets with a thickness of 0.8 mm, 1.0 mm and 1.2 mm. The most loaded is the lining of the ceiling panel between frames No. 7 and 13, where the thickness of the lining is 1.2 mm. The lining of the left panel of the superstructure in the area between frames No. 19 and 23 is made of a sheet 1 mm thick.

Cargo wings (Fig. 2.12) are located between frames No. 13 and 21 of the central part of the fuselage, each is suspended on two loops to an inclined frame.

Load flaps close the rear opening in cargo compartment and create additional cabin volume. Doors of riveted design, each consists of stamped stiffness and outer duralumin cladding. For the convenience of loading wheeled vehicles, the sashes have flaps 13 that fold upwards, which are hinged to the lower parts of the sashes. In the tilted position, the shields are held by rubber shock absorbers.

Opening and closing of the cargo flaps is done manually, in the open position they are held by struts, and in the closed position they are fixed with pins at frame No. 13 and locked with longitudinal and transverse locks 10 and 11. The locks allow opening the flaps from inside the cargo compartment.

Rice. 2.10. Ceiling panel:

1 - engine mounts; 2,3 - attachment points of the engine fixing device; 4 - attachment point of the struts of frame No. 1, hood and fan; 5 - attachment points of frame No. 1 of the hood; 6 - a hatch to the booster pumps of the supply tank; a - holes for the bolts of the frame of the main gearbox.

Rubber profiles are reinforced on the end surfaces of the wings along the entire perimeter, which ensure the sealing of the mating surfaces of the wings with the fuselage and between themselves in the closed position. To exclude the opening of the cargo doors when the helicopter is parked outside, a fixing device for the inner handle of the door lock is installed; before departure, the handle must be unlocked.

Toolboxes 12 are installed in the lower part of the wings. Both doors have hatches for exhaust gases from the running engine of the transported equipment in the cargo compartment. On the left wing there is a portable fire extinguisher 16 and brackets for fastening the lodgements under the racks 17 of the sanitary stretcher. In the outer skin, hatches are cut out under the blinds with an exhaust ventilation damper 1 and under the rocket launchers 2. On the right wing there is a hatch closed by a lid for supplying the sleeve of the ground heater 6.

The right wing is equipped with a hatch for leaving the helicopter in an emergency. The hatch is closed with cover 8, which consists of outer skin and rigidity riveted together. At the bottom, the manhole cover is held by latches, and at the top - by locking pins of the emergency drop mechanism mounted on the cover.

The emergency ejection mechanism is similar in design to the cockpit sliding blister mechanism. To drop the lid, you need to sharply pull the handle 7 down, then the locking pins will come out of the lugs of the brackets and release the lid, and the spring pushers located in the upper corners of the hatch will push the lid out.

Ladders 15 are attached to the helicopter, designed for loading and unloading wheeled vehicles and other cargo. In the working position, the ladders are fixed with steel knots in steel sockets on the lower beam of frame No. 13, in the stowed position they are laid and fixed on the floor on both sides of the cargo compartment. Depending on the load of the helicopter, if it is impossible to place cargo ladders on the cabin floor, the ladders are placed on the left wing of the cargo hatch, where ladder attachment points are provided in the stowed position.

Rice. 2.12. Load doors:

1 - damper for exhaust ventilation; 2 - rocket launcher; 3 - folding seat; 4 - the door of the boar crew; 5 - electric winch; 6 - hatch for supplying the sleeve of the ground heater; 7 - reset handle emergency hatch cover; 8 - emergency hatch covers; 9 - handle; 10 - pin lock; 11- coupler lock; 12 - tool box; 13 - shield; 14 - seat; 15 - ladders; 16 - portable fire extinguisher; 17 - mounting bracket for sanitary racks.

The frame of the gangway consists of a longitudinal and transverse power set. The longitudinal power set consists of two beams riveted from corner profiles and D16T L1, 2 duralumin wall. The upper chords of the beams are made of D16T duralumin T-section, the shelf of which protrudes above the ladder sheathing and prevents wheeled vehicles from rolling off the ladders during its loading and unloading. The transverse set consists of T-profiles and stamped diaphragms made of duralumin sheet riveted to them.

The front and rear edges of the ladders have steel edging. To prevent slipping of the wheels of self-propelled equipment when loading it under its own power, corrugated linings are riveted to the edgings on the rear end parts of the ladders.

Rice. 2.11. Covering the central part of the fuselage

4. TAIL BOOM

The tail boom provides the creation of the shoulder necessary for the tail rotor thrust to compensate for the reactive moment of the main rotor.

The tail boom (Fig. 2.14) of riveted construction, beam-stringer type, has the shape of a truncated cone, consists of a frame and smooth working duralumin skin.

The frame includes longitudinal and transverse power sets. The transverse power set consists of seventeen Z-section frames. Frames No. 1 and 17 are docking, they are made of extruded D16AT duralumin profile and reinforced with toothed bands. Frames No. 2, 6, 10 and 14 are reinforced in the upper part for supports 3 of the transmission tail shaft. Brackets 2 are also attached to them for installing textolite guide blocks for tail rotor pitch control cables.

The longitudinal set consists of 26 stringers #1 through #14, starting at the top on either side of the vertical axis. Stringers are made of extruded angle profiles.

The tail boom skin is made of D16AT sheet clad duralumin. The joints of the skin sheets are made along the stringers and frames with an overlap with undercut. In the skin between frames No. 13 and 14, on both sides of the tail boom, cutouts were made for the passage of the stabilizer spar.

Rice. 2.14. Tail Boom:

1 - docking flange; 2 - bracket for fastening the blocks of tail rotor control cables; 3 - transmission tail shaft support; 4 - adjusting bracket assembly; 5 - overlay; 6 - stabilizer mounting bracket; 7 - attachment point of the shock absorber of the tail support; 8 - attachment points of the tail support strut.

Reinforcing duralumin plates 5 are riveted along the contour of the cutouts. On top of the skin there are hatches with covers for inspecting and lubricating splined couplings of the transmission tail shaft. Between frames No. 3 and 4, a cutout was made for the MSL-3 flashing beacon, between frames No. 7 and 8, 15 and 16 - cutouts for drill lights, between frames No. 11 and 12 - a cutout for the course system sensor.

From the bottom of the tail boom between frames No. 1 and 6, a radome for the antenna of the DIV-1 device is installed. The upper part of the fairing is riveted from duralumin profiles and skin, fastened to the beam with screws. The lower part is made of a radio-transparent material, fixed to the upper part on a ramrod rod and is locked with two folding locks and three plates with screws. Two antennas (receiving and transmitting) of the RV-3 radio altimeter are installed on the lower part of the beam. On frame No. 13 on both sides of the beam, nodes 4 are installed for the bolts of the stabilizer adjusting brackets, and on frame No. 14 - brackets 6 for mounting the stabilizer. On frame No. 15, on both sides of the tail boom, attachment points 8 for the tail support struts are riveted, and on frame No. 17 from the bottom - assembly 7 for attaching the tail support shock absorber.

5. END BEAM

The end beam (Fig. 2.15) is designed to move the axis of rotation of the tail rotor into the plane of rotation of the main rotor in order to ensure the balance of the moments of forces relative to the longitudinal axis of the helicopter.

Rice. 2.15. End beam:

1 - frame No. 3; 2 - frame No. 9; 3 - fixed part of the fairing; 4 - wall of the spar; 5 - tail light; 6 - inclined antenna; 7 - removable part of the fairing; 8 - cover; 9 - keel beam.

The riveted end beam consists of a keel beam 9 and a fairing. At frame No. 2, the axis of the beam has a break at an angle of 43 ° 10 "in relation to the axis of the tail boom.

The frame of the keel beam consists of a transverse and longitudinal set. The transverse set includes nine frames. Frames No. 2, 3 and 9 are reinforced, and frame No. 1 is docking.

The longitudinal set consists of a spar 4 and stringers made of corner profiles. The spar of riveted design is made of D16T duralumin corner profiles, the walls are made of duralumin sheet. In the lower part of the wall of the spar there is a hatch for access to the intermediate gearbox. The frame of the keel beam is sheathed with a smooth running sheathing made of D16AT duralumin, on the right side 1 mm thick, on the left - 1.2 mm. Between frames No. 1 and 3, a reinforced sheathing made of D16AT duralumin 3 mm thick is installed, on the inside of which, to facilitate, longitudinal milling was made, made by a chemical method. A similar skin 2 mm thick is riveted between frames No. 8 and 9.

Docking frame No. 1 is stamped from aluminum alloy D16T, to increase the reliability of the joint, the thickness of the joined planes is increased to 7.5 mm with their subsequent machining.

Reinforced frame No. 3 (pos. 1) is a bracket stamped from AK6 aluminum alloy, an intermediate gearbox is attached to it with four bolts, and a tail gearbox is attached to the flange of frame No. 9. There are two hatches at the top of the beam bend - upper and lower. The upper hatch is designed for filling oil into the intermediate gearbox, and the lower hatch is for inspecting the spline connection. The hatches are closed with lids, which have gill slits for air intake for cooling the intermediate gearbox. During operation, both hatches are used to install a fixture when measuring the angle of fracture between the tail and end shafts of the transmission.

The fairing forms the rear contour of the keel beam and is a fixed rudder that improves the directional stability of the helicopter. The fairing is made of two parts - the lower 7 is removable and the upper 3 is non-removable. The fairing frame consists of six stamped stringers made of D16AT duralumin, six ribs and docking tapes riveted along the contour of the fairing.

The frame is sheathed with smooth duralumin sheathing. In the lower part of the fairing there is a hatch, in the cover 8 of which gill slits are made for the exit of air cooling the intermediate gearbox. In addition, inclined antennas 6 are mounted on both sides, and whip antennas are mounted along the axis of symmetry of the fairing. A tail light is installed behind the axis of symmetry of the fairing. The removable part of the fairing is fastened to the belts of the spar of the keel beam with self-locking screws, and the fixed part - with rivets using butt bands.

Fig.2.16. Scheme of docking the fuselage with a typical

connection of docking frames (below)

The docking of the fuselage parts is of the same type and is carried out along the docking frames in accordance with the scheme (Fig. 2.16). All docking frames are made of extruded D16AT duralumin profile, the end shelf of which forms a flange with holes for docking bolts.

To reduce the stress concentration in the skin along the contour of the docking frames, duralumin toothed tapes are laid, which are riveted together with the skin to the outer flange of the frame.

6. STABILIZER

The stabilizer is designed to improve the characteristics of the longitudinal stability and controllability of the helicopter. The stabilizer (Fig. 2.17) is installed on the tail boom between frames No. 13 and 14, its mounting angle can only be changed when the helicopter is on the ground.

The stabilizer has a NACA-0012 symmetrical profile and consists of two halves - right and left, symmetrically located relative to the tail boom and interconnected inside the boom.

Both halves of the stabilizer are similar in design. Each half of the riveted stabilizer consists of a spar 2, seven ribs 5, a tail stringer 12, a diaphragm, a frontal duralumin sheathing 6, a removable end fairing 9 and a fabric sheathing 11.

The ribs and diaphragms are stamped from sheet duralumin. The ribs have nose and tail parts, which are riveted to the spar belts. Ridges with holes for sewing on linen sheathing are made on the shelves of the tail parts of the ribs.

The tail stringer, made of sheet duralumin, covers the tails of the ribs from below and above and forms a rigid trailing edge of the stabilizer. The tails of the ribs with a tail stringer are riveted with flush rivets.

Rice. 2.17. Stabilizer:

1 - stabilizer linkage axis; 2 - spar; 3 - adjusting bracket; 4 - docking flange; 5 - rib; 6 - duralumin sheathing; 7 - beam antenna attachment point; 8 - balancing weight; 9 - end fairing; 10 - drainage hole; 11 - linen sheathing; 12 - tail stringer.

On the toe of rib No. 1 of each half of the stabilizer, a bracket 3 with an earring is riveted, with which you can change the installation angle of the stabilizer on the ground.

A balancing weight 8 weighing 0.2 kg is riveted to the front of rib No. 7, covered with a removable end fairing 9 made of fiberglass. On the toe of rib No. 7 of the right and left halves of the stabilizer, node 7 is installed for attaching the cord of the beam antenna.

Stabilizer spar of beam type of riveted structure consists of upper and lower chords and a web with beaded holes for rigidity. The upper and lower belts of the spar are made of duralumin corner profiles. In the root part, the spar is reinforced with an overlay riveted to the belts and the spar wall from the rear side, and in the front part between ribs No. 1 and 2, the spar is reinforced with an overlay riveted to its belts. A docking flange 4, stamped from an aluminum alloy, is riveted to the overlay.

Fittings with axles 1 are installed on the spar near rib No. 1 for hanging halves of the stabilizer on the tail boom. The stabilizer attachment points are protected from dust by covers, which are fastened to the spar and rib No. 1 with a cord and a clamp using a foam plastic boss.

The nose of the stabilizer is sheathed with D16AT duralumin sheets riveted along the shelves of the bow parts of the ribs and the spar belts. The tail section is sheathed with AM-100-OP fabric, the seams along the ribs are sealed with toothed tapes.

Docking of the right and left halves of the stabilizer is made by bolts on the docking flanges and connecting plates.

The fuselage of a helicopter is the body of an aircraft. The helicopter fuselage is designed to accommodate the crew, equipment and payload. The fuselage can accommodate fuel, landing gear, engines.

In the process of developing the volume and weight layout of the helicopter, the fuselage configuration and its geometric parameters, coordinates, magnitude and nature of the loads that must be perceived by the power elements are determined. The choice of the fuselage KSS is the initial stage of design. Such a power scheme is being worked out, which most fully meets the requirements of the customer.

Basic requirements for the fuselage KSS:

reliability of the design during the operation of the helicopter;

ensuring a given level of comfort in the cockpits of the crew and passengers;

high operational efficiency;

ensuring a safe volume for the crew and passengers inside the fuselage and the possibility of leaving it during an emergency landing of a helicopter.

The operational requirements, design and purpose of the helicopter also significantly affect the choice of the fuselage KSS. These requirements are:

• - maximum use of the internal volumes of the fuselage;
• - providing the visibility required for the helicopter crew;
• - providing access for inspection and maintenance of all units located in the fuselage;
• - convenient placement of equipment and cargo;
• - convenience of loading, unloading, fixing cargo in the cab;
• - ease of repair;
• - sound insulation, ventilation and heating of the premises for passengers and crew;
• - possibility of replacement of glasses of a cabin in operating conditions;
• - the possibility of re-equipment of passenger cabins by changing the layout of the room, the type of seats and the step of their installation.

For emergency escape of the helicopter by passengers and crew, emergency exits are provided on the helicopter. Doors for passengers and crew, as well as maintenance hatches are switched on

in the number of emergency exits, if their size and location meet the relevant requirements. Emergency exits in the cockpit are located one on each side of the fuselage, or instead there is one upper hatch and one emergency exit on either side. Their size and location should ensure that the crew can quickly leave the helicopter. Such exits may not be provided if the helicopter crew can use the emergency exits for passengers located near the cockpit. Emergency exits for passengers should be rectangular in shape with a corner radius of not more than 0.1 m.

The dimensions of the emergency exits for the crew must be at least:

480 x 510 mm - for side exits;

500 x 510 mm - for a rectangular top hatch or G40 mm in diameter - for a round hatch.

Each main and emergency exit must meet the following requirements:

Have a movable door or a removable hatch that provides free exit for passengers and crew;

Easy to open both from inside and outside with no more than two handles;

Have means for locking the outside and inside, as well as a safety device that prevents the opening of the door or hatch in flight as a result of accidental actions. Locking devices are made self-locking, without removable handles and keys. On the outside of the helicopter, places are indicated for cutting down the skin in case of jamming of doors and hatches during an emergency landing of the helicopter.

The volumes required to accommodate passengers and the transported cargo are decisive in the design of the passenger and cargo cabins of the fuselage.

The appearance of the fuselage and its KOS depend on the purpose of the helicopter and its layout:

The amphibious helicopter must have a special shape of the lower part of the fuselage that meets the requirements of hydrodynamics (minimum loads on the helicopter when landing on water; the minimum required thrust is 11V during takeoff; no splashing in the pilot’s field of vision and engine air intakes; compliance with the requirements of stability and buoyancy );

The crane helicopter fuselage is a power beam to which the cockpit is attached, and the cargo is transported on an external sling or in containers connected to the butt joints of the lower central part of the fuselage;

In the most common single-rotor helicopter scheme, it is necessary to have a power cantilever beam for attaching the RV.

The choice of a rational CSS of the fuselage is carried out primarily on the basis of weight statistics, parametric dependencies and generalized information about the power circuits of previous designs.

Based on the results of the decisions made, proposals are formed, on the basis of which the final choice of the fuselage KSS is made. In most cases, based on the requirements and operating conditions, it is already known in advance which type of design is applicable in one case or another, so the task can be reduced to finding the best option within a given design type.

In frame structures, KSS, which have already been proven by long-term practice, are used - these are structures such as reinforced shells (beam scheme), truss structures and their combinations.

The most common beam scheme of the fuselage. The main reason for the development of beam fuselages is the desire of the designer to create a strong and rigid structure in which the material, optimally distributed over a given section perimeter, is rationally applied under various loads. In the beam structure, the internal volume of the fuselage is used to the maximum, all the requirements of aerodynamics and technology are met. Cutouts in the skin require local force, which increases the mass of the fuselage.

Beam fuselages are divided into two types - spar and monoblock.

The scheme of the fuselage is significantly modified in the presence of cutouts in the design, especially at their considerable length. As the sections approach the end part of the cutout, the stresses in the skin and stringers are significantly reduced, the transmission of torque becomes more complicated, and additional stresses appear in the longitudinal set. To maintain the strength of the panel, the stringers along the cutout boundary are reinforced, turning into spars. Sheathing and stringers are fully included in the work only in the section located from the ends of the cutout at a distance equal to approximately the width of the cutout. In such a case, it is advisable to take the KSS of the fuselage as a spar.

In spar structures, the bending moment is perceived mainly by longitudinal elements - spars, and the skin perceives local loads, shear force and torque.

In a monoblock design, the skin, together with the frame elements, also perceives normal forces from bending moments.

A combination of these power circuits are stringer fuselages with partially working skin, which is made in the form of a thin-walled shell, reinforced with stringers and frames. A variety of monoblock KSS is.

Monocoque made of homogeneous material. It provides for the presence of only two elements - skin and frames. All forces and moments are perceived by the skin. This scheme is most often used for tail booms of small diameters - D< 400 мм (обшивка, согнутая по цилиндру с малым радиусом, имеет высокую устойчивость при сжатии).

The monocoque is multi-layered. The use of three-layer panels with thin bearing layers makes it possible to increase both local and general rigidity of fuselage parts with a regular (without cutouts) zone. The design of three-layer (layered) panels is very diverse and depends on the materials of the outer and inner layers, the type of filler, the method of joining the skins with the filler, etc.

The surface of the fuselage, used for moving technical personnel during ground handling of the respective units, is made of panels of a layered structure (increased rigidity) with a thickened outer bearing layer with a friction coating. These panels must be included and the power circuit of the fuselage.

It is advisable to perceive the load from soft fuel tanks with panels of a layered structure. These panels, having high bending rigidity, simultaneously serve as a tank container, and then it is not necessary to create an additional bearing surface based on the stringer set of the lower fuselage.

KM has been successfully introduced into the design of the helicopter airframe, and has already been operated on several generations of helicopters.

Modern glass-reinforced plastics can compete with traditional aluminum alloys in terms of specific strength, but they are significantly, at least 30% inferior to them in terms of specific rigidity. This circumstance was a brake on the way to expand the use of fiberglass and structural elements.

Organoplastics - materials that are lighter than fiberglass in terms of specific rigidity are not inferior to aluminum alloys, and in terms of specific strength they are 3-4 times superior. The wide development of organoplastics made it possible to set a fundamentally new task - to move from the creation of individual parts from CM for metal structures to the creation of the structure itself from CM, to their extended use, and in some cases to the creation of a structure with the predominant use of CM.

KM are used both in the skins of three-layer panels of tail, wing, fuselage, and in frame parts.

The use of organic instead of fiberglass allows to reduce the weight of the airframe. In heavily loaded units, organoplastics can be most effectively used in combination with other more rigid materials, such as carbon fiber.

Structural and technological scheme of the fuselage of an experimental Boeing-360 helicopter, all the power elements of which are made of panels of a layered design using a composite material.

The use of thin skins, well reinforced with honeycomb filler (having a low density), makes layered structures a reserve for reducing the weight of the fuselage. High specific strength and resistance to vibration and acoustic loads determine the growth in the use of such structures as load-bearing elements of the fuselage.

The potential advantages of three-layer structures can only be realized if production is organized at a high level. technical level. The issues of design, strength and technology of these structures are so closely interconnected that the designer cannot but pay attention to great attention technological issues.

The long-term strength of glued joints and the tightness of honeycomb units (from moisture ingress) are the main things that should be ensured by structural and technological development.

Technological issues include:

• - choice of brand of glue that provides the necessary strength with an acceptable weight gain;
• - the ability to control technological regimes at all stages of manufacturing units;
• - providing a given degree of coincidence of the contours of the mating parts (mainly honeycomb block and frame);
• - application of reliable methods of control with measuring the strength of gluing;
• - choice of additional sealing method;
• - the introduction of honeycombs without perforation.

Farm fuselage. In the fuselage of the truss scheme, the power elements are spars (truss belts), racks and braces in the vertical and horizontal planes. The skin perceives external aerodynamic loads and transfers them to the truss. The farm perceives all types of loads: bending and torsional moments and shear forces. Due to the fact that the skin is not included in the power circuit of the fuselage, the cutouts in it do not require significant reinforcements. The presence of rods in the truss structure makes it difficult to use the internal volume of the fuselage, the placement of units and equipment, their installation and dismantling.

Elimination of resonant vibrations of numerous rods is a difficult task. The truss structure makes it difficult to meet the aerodynamic requirements for the shape of the fuselage and the rigidity of the skin. In this design, it is difficult to apply advanced technology for welding assemblies with a complex configuration. weld. Heat treatment of a large truss after welding is associated with certain problems. The listed main disadvantages of the truss structure are the reason for their limited.

The CCC of the cabin floor is determined by the purpose of the helicopter. In a transport helicopter for transporting wheeled vehicles, the cargo floor must be reinforced with longitudinal beams placed in such a way that the loads from the wheels are perceived directly by these load-bearing elements. To fix wheeled vehicles in the floor, knots are installed for fastening bracing cables at the intersection of the longitudinal (stringer) and transverse (frame) frame elements. Monorails mounted on the cabin ceiling are used for loading and unloading containers. The load on the cables is attached to the trolley, fixed to the monorail, and moves along it to a predetermined place in the cabin. It is advisable to include monorails in the power circuit of the fuselage. In the cargo compartment, mooring knots are also installed with the required interval for the corresponding loads.

For the convenience of loading and unloading oversized cargo, the cargo ladder (ramp) should be mechanized so that it can stop and lock in any position, as well as to ensure the possibility of transporting goods on an open rear ladder.

The power elements of the fuselage are mainly made of aluminum alloys. In places exposed to heat, titanium is used and stainless steel. The fairings of the power plant and tail transmission (located on top of the tail boom) are rationally made of fiberglass reinforced with reinforced stiffeners.

When forming the KSS of the frame unit, it is necessary to take into account the following main provisions:

The distance between the power transverse elements and their placement on the unit is determined by the place of application of concentrated forces normal to the axis of the unit;

All concentrated forces applied to the frame elements must be transferred and distributed to the skin, through which they are usually balanced by other forces;

Concentrated forces must be perceived by frame elements directed parallel to the force through stringers and spars, and forces acting across these units, by frames or ribs, respectively;

Concentrated forces directed at an angle to the axis of the unit must be transmitted to the skin through the longitudinal and transverse load-bearing elements. The force vector must pass through the intersection point of the stiffness axes of these elements;

The cutouts in the frame unit must have expansion joints along their perimeter in the form of reinforced belts of longitudinal and transverse elements.

The presence of cutouts in the power structure of the fuselage, abrupt transitions from one configuration to another, and zones of application of large concentrated forces (i.e., “irregular zones”) have a significant effect on the distribution and nature of the force flow of stresses, which is similar to the fluid velocity field in the region of local resistances.

The stress concentration in the fuselage structural elements, the amplitude and frequency of alternating stresses are the determining parameters in solving very important issue creation of a high-resource fuselage.

The problem associated with the design of the fuselage can be solved in the following ways:

Develop the CSS taking into account the analysis of the nature and place of application of external forces and operational requirements that determine all kinds of cutouts (their size, location on the fuselage);

Use thin (no moment) sheathing, which can lose stability under short-term high loads without permanent deformation;

On the basis of sufficient experience in production and operation, widely introduce elements made of CM into the practice of designing frame units.

The final formation of the fuselage CSS of the minimum mass with a given resource is carried out on the basis of an analysis of the results of experimental studies of a full-scale frame for calculated cases of loading of power elements with a complete simulation of the forces and moments applied to the fuselage.