Raduga KSR-5 (AS-6 Kingfish)
According to the aerodynamic scheme the KSR-5 was a monoplane with a mid-wing and plumage. The KSK-5 missile has an all-metal construction with a working lining, a mid-planimetric arrangement of a small triangular wing and a cross-shaped plumage. The heel and pitch control was carried out by an all-terrain stabilizer, working in elevon mode, at the rate - all-terrain upper keel. To increase flight stability, the missile had a sub-body bottom keel, which folded to the right in flight during storage, transportation and suspension. The launch arrangement was provided by a pneumatic cylinder together with the rudders' unfolding by the firing clamp.
The main structural materials of the airframe of the rocket were stainless steel EI-654 (oxidizer tank), 30???? and 12?2???? (power units of plumage and butt elements), alloys ???-6? (fuel tank) and ?16? (part of the shells and panels). Heat-resistant steel 1 2????? was also used to make a sock of the rocket fairing that was heated in flight to 420°. Large hull units were all-welded, wing and plumage were made with a wide use of honeycomb panels made of thin aluminum foil, which for strength and rigidity were poured molten xylite, and then processed along the theoretical contour on a vertical milling machine using a copier. Large-size coloured castings were widely used for the production of power frames and equipment fastening beams.
A number of problems were caused by the manufacture of radio transparent fairings - large-size products, which at high supersonic aerodynamic and thermal loads should have the necessary mechanical strength, thermal and moisture resistance, high surface purity, low weight and, most importantly, the proper radio transparency, which was subject to particularly stringent requirements. The latter had a direct impact on the characteristics of CNS, which required a radio transmission coefficient of 70-75%. Experiments and searches for the technology of manufacturing the lightest in the required radio range of the fairings, carried out by leading engineers of plant # 256 V.N.Lezhenin, L.E.Kuznetsova, A.S.Kazakov and specialist VIAM K.T.Scherbakova were required.
The practice of manufacturing solid fairings for X-22 from glass fiberglass has shown that they have too much weight. The fairing of almost two meters in size with the required strength and rigidity should have a small weight, but precisely defined contours. For radio-transparency requirements, wall thicknesses were limited to 4-7 mm. The cone had a honeycomb structure with mesh filler, and the thickness varied from larger at the toe to thin at the base. For the production of its outer and inner jackets steel punches were used, on which fiberglass coverings were molded under vacuum. On the same punch, a mesh mesh mesh was glued to the inner shirt, molded along the contour, and then the outer shirt and power belt were put on at the base. The final polymerization of the bundle with curing of the binder resin for strength was carried out in the furnace in accordance with the stepped temperature regime.
The problem was the design of the balloons that fed the missile's steering drives and the boost system. The balloons were made of two thin-walled 5-mm thick chromancer hemispheres, which were welded together. Cylinders had to withstand a working pressure of 350 atm, but the products were not always durable. In operation, there were a number of cases of their explosion, which destroyed the structure, with the engine and tanks nearby, which led to a fatal outcome for the entire product and was extremely dangerous for personnel. Cylinders - "bombs" were not trusted until they were reinforced by winding fiberglass bundles on the body. This design withstood pressure of up to 700 atm, allowing to reduce the thickness of the walls to 4 mm, and at the rupture did not give splinters, splitting into halves.
The nose section of the missile was occupied by the VS-KN homing equipment with active radar CNS. With the help of aircraft target designation CNS captured the target on the suspension, tracking its position on the azimuth. The control system - "take-off" provided homing on course and pitch and software control of altitude, which provides the output of the missile in the stratosphere with a subsequent dive on the target. Later on, the control system was improved to provide a counter-air maneuver in case of enemy radar irradiation.
The next compartment accommodated a 700 kg BC of blast-current type or a nuclear BC in a special container with all necessary equipment, including an explosive device that provides a specified ground or air blast, a safety and launch system with detachment sensors from the carrier and trajectory sensors, a temperature control system that maintains the necessary temperature and humidity in the compartment.
The central part of the missile was occupied by a fuel cell with fuel and oxidizer tanks. The fuel tank contained 660 liters of fuel and the steel oxidizer tank contained 1010 liters of oxidizer. The pressure of the tanks was provided by an air system that simultaneously served as the first stage of the feed system, it also used to inflate the waveguides and arrange the lower keel. The pipes of the inflatable system, feeder and wiring of the electric fittings were laid in the basement gargota under removable panels to provide access. "Dry" ampoule batteries with converter in the equipment compartment provided power supply to the systems for 480 sec. The low-altitude KSR-5H had a number of differences in the power system and storage batteries.
The missile was equipped with a C.5.33 type LCD designed by A.M.Isayev Design Bureau with two combustion chambers with separate exhaust nozzles. The supersonic nozzles are unregulated and each of them was optimized to create a certain thrust. The engine was powered by a common turbo-pump unit (TNA) with automatic mode adjustment, performing two programs with different traction. The high-performance TPA provided the required fuel consumption of 80 kg/sec and the required pressure at the engine inlet (the LPG thrust increases significantly as the operating pressure in the combustion chamber increases, and the feed pressure must exceed it to allow fuel to enter the engine). The dual-chamber design of the LRD, with its compact size and light weight, provided the required range of thrust required in various flight modes. When the rocket was launched, a combustion chamber began to operate, the thrust of which in 7100 kgf reported to the rocket thrust-to-weight ratio of 1.8 and provided a fast acceleration and altitude gain. To maintain the speed at altitude, a marching combustion chamber with an economical thrust of 600 kgf or 1120 kgf, depending on the set flight mode, was included in the work. The fuel was two-component: fuel TG-02 (660 l) and oxidizer AK-27P (1010 l). The air system of the missile provided the waveguides, oxidizer and fuel tanks, lower keel layout and other operations.
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