- combat unit
- radio detonator
- air pressure accumulator
- fuel tanks
- swivel wing
- steering wheel
- radio control equipment
- isopropyl nitrate tank
- starting accelerator
- turbine pumping unit
- nozzle block
- combustion stabilizer
The main developer of the ZMD (designated as ZM8) - OKB-8 - was unequivocally given the use of a direct-flow air jet engine (DFID) on an anti-aircraft guided missile. The choice of this type of engine using non-aggressive liquid fuel seemed to be quite reasonable. Air oxygen was used as an oxidizer in the RVRP, so that the rocket tanks contained only fuel - kerosene. The SARD outperformed rocket engines in specific thrust five or more times. For missile flight speeds 3-5 times higher than sound, the SARD had the lowest fuel consumption per thrust unit, even compared to a turbojet engine. Compared to it, the direct-flow engine design seemed surprisingly simple, and was much cheaper. Almost the only drawback of the SARD was considered to be the inability to create significant thrust at subsonic speeds in the absence of the necessary velocity pressure at the intake, which did not allow limiting the use of only SARD on missiles launched from the Earth.
By the end of the 1950s, the greatest experience in the design and development of SARD was gained at Design Bureau-670 of the State Committee for Aircraft Engineering (SCAT) by the team headed by chief designer M.M.Bondaryuk. Their most significant work was the creation of a supersonic SARD for S.A.Lavochkin's intercontinental cruise missile "Storm", which was successfully tested both on the stands and in flight tests. The engines for V.M.Myasishchev's similar rocket "Buran" as well as for other aircraft were under development. However, the existing experience was somewhat one-sided - the engines were developed for small maneuverable vehicles flying at a constant speed at almost the same altitude. Taking into account the impossibility of SARP operation at low speeds, the 3M8 rocket was made according to a two-stage scheme. To ensure the conditions for launching a direct-flow engine, solid-fuel accelerators accelerated the rocket to a speed corresponding to the number M=1.5 ... 2.
By the end of the fifties already had information about the unstable nature of direct-flow engines at large angles of attack. At the same time, for the anti-aircraft missile intended for defeat of highly maneuverable planes of front-line aviation, realization of transverse overloads of an order of eight units was required. This largely determined the choice of the general scheme of the missile. For the second (marching) stage the layout with a turning wing which provided possibility of creation of sufficient lifting force at small angles of attack of the rocket case has been accepted. The body of the missile's marching stage was a supersonic direct-flow engine 3C4 - a pipe with a pointed central body, ring nozzles and stabilizers for combustion. On previous missiles of similar schemes most of the systems and units were placed on a circular scheme in the outer casing of the SARD. However, this arrangement was clearly contraindicated to a number of elements, such as a combat unit. In the central body of the air intake with a 450 mm cylindrical part diameter, in addition to a 3H11 fragmentation and high-explosive warhead weighing about 150 kg, there was a 3E26 radio detonator and an air pressure accumulator balloon. A homing head was supposed to be installed in the front part of the central body. The central body was slightly buried in the inner volume of the missile body. Then there were openwork constructions from ring and radial elements - straightening grids, nozzle blocks, combustion stabilizers. In the ring casing of the engine with an external diameter of 850 mm starting from its front edge were located tanks with kerosene, approximately in the middle of the length - steering machines, wings, and closer to the rear edge - blocks of control system equipment (SC).
Swivel wings of 2206 mm in scope were placed according to "X" pattern and could be deviated by hydropneumatic steering in the range of +/-28°. The chord of the wing was 840 mm at the base, 500 mm at the end. Arrowhead on the front edge was +19 ° 38 ', on the rear edge of -8 ° 26 ', the total area of both consoles (swivel parts) in one plane - 0.904 square meters. m. Stabilizers with a swing of 2702 mm were installed according to "+" - figurative scheme. The chord is 860 mm at the base, 490 mm at the end. The front edge - with an arrowhead of 20 °, the rear edge - straight, the total area of two consoles in one plane - 1.22 square meters. m. The length of the missile was 8436 mm, diameter 850 mm.
At starting weight of 2455 kg the initial weight of the second (marching) stage was about 1400 kg, from which about 270 kg were fuel - kerosene T-1 (or TC) and 27 kg for isopropyl nitrate. Fuel supply was provided by the turbo-pump unit C5.15 (on the first samples - C2.727), working on mono fuel - isopropyl nitrate. This unitary fuel, in comparison with hydrogen peroxide, which was widely used in rocket engineering earlier, at somewhat lower density (about a quarter) had bigger power and, what is more important, was more stable and safer in operation.
Each of the four launch engines 3C5 was equipped with a charge 4L11 solid ballistic fuel RSI-12K weighing 173 kg in the form of a single-channel draughtsman length of 2635 mm with an outer diameter of 248 mm and a channel diameter of 85 mm. To ensure separation of the starting engines from the marching stage, a pair of small aerodynamic surfaces located at an angle to the longitudinal axis of the engine were fixed on each of them in the stern and nose parts.
The 3M8 rocket initially provided for the application of combined control - a radio command system in the main flight and homing on the final section of the ZUR trajectory. Semi-active radar homing head was to operate on the reflected from the target signal pulse radiation channel target tracking station guidance missiles.