RT-20P / SS-15 SCROOGE - Design
The RT-20P had a very un-American design - the missile had a storable liquid-propellant second stage, and a solid propellant first stage.
The UK Ministry of Defence, Directorate of Scientific And Technical Intelligence (DSTI), provided an assessment of the SCROOGE in December 1966 after the Moscow parade of November 1966. The tracked prime mover was similar to that of SCAMP and was a development of the JOSEPH STALIN [JS] chassis used for heavy tanks in the Great Patriotic War. There were however, slight differences between the SCAMP and SCROOGE chassis. SCROOGE had eight road wheels instead of six, reinforced shock absorvber and re-arranged top rollers. The prime mover carrying capacity was assessed as approximately 150,000 pounds.
SCROOGE body work was very different from SCAMP. The canister [which can be elevated to the vertical] is supported by a framework extending from the fron tof th body; two hydraulically-operated legs at the rear support the vehicle when the canister is elevated. The pivot point from the canister can be seen quite clearly and its position permits ground clearance for the canister when elevated. Three clamps secrue the canister to the prime mover. A tubular framework supports it while two hydraulic jacks elevate it to the vertical.
The canister was of welded construction, apparently stiffened internally. There was no sign of a longitudinal break or hinge as on SCAMP. Two conduits extended amost the complete length of the canister. The one on the port side extended to a cylindrical casing projecting from the rear of the canister while the other counduit appeared to be connected to a tubular housing projecting from the conical nose. Twenty service connections leading form the primve mover fed the starboard conduit through a junction box. The rear canister face was a conical frustrum with a cylindrical casing about 2 feet in diameter bolted to it. There appeared to be no provision for remoing the rear face from the canister tube.
The design and construction of the canister implied that it would most probably act as a launching tube as well as an environmentally controlled transporter. The absence of any apparent means to allow the escape of the rocket efflux at the rear suggested that the missile was ejected before ignition takes place [cold launch technique].
The cylindrical housing on the rear face could well hold a gas generator for this purpose. The circular hatch is probably positioned to permit access to the missile guidance bay. With the possible exception of the re-entry vehicle, this was the only part of the missile which would need setting-up by some externatl means after it had been loaded into the canister and positioned at its firing point.
The housing at the nose probably contained an explosive device for removing the end-cap just before launch. Two large brackets, one on each side of the canister in the plane of the front mounts apparently had no function when the launch tube was horizontal. However, they would provide good lateral support points when the cannister was elevated.
The fuel compartment of the II stage is welded in the form of a single tank, separated by an intermediate bottom of the waffle structure on the cavities "O" and "G". To reduce the dynamic loading of structural elements due to fluid oscillations that arise during the transportation of the loaded rocket, the dams of the fuel compartment use spherical-shaped dampers that maximally approximate the "tank-liquid" system to the solid body. To ensure the necessary conditions for transportation and launching the missile, fuel components filled with gases to equilibrium concentrations are filled in fuel cavities.
Aluminum alloys are used as structural materials. The fuel compartment is equipped with counter-pressure nozzles that use the gas to pressurize the cavity "O" for braking the II stage while separating the BB. The single-stage II stage engine is mounted on a transitional welded frame fixed to the rear end frame of the fuel compartment. On the back of the fuel compartment LLC containers are mounted, ensuring the shooting of the latter with the required speeds and direction to ensure the construction of an effective combat chain in cooperation with the BB.
The I stage adapter is a cylindrical riveted compartment consisting of a skin, a power set, a stabilizing shield and a protective cone. The lower base of the adapter is attached to the solid-propellant engine of the 1st stage, and the upper one - to the discontinuous
bolted to the fuel compartment of the II stage. The cover of the adapter is made of aluminum alloy D19AT. In the lower part of the cladding, 16 windows with a total area of ??1.2 m2 are cut out for the exit of the hot gases of the engine of stage II during the gas dynamic separation of the stages.
Inside the adapter is attached a stabilizing screen, made in the form of a truncated cone. To prevent the reflected gas-dynamic jet from impacting the control and telemetry devices located under the screen, they are covered with a reverse protective cone of a similar design. Stabilizing screen in the zone of action of the gas-dynamic jet of the engine of the II stage is covered with a heat-protective cover made of asbestos cloth.
The engine body of the 1st stage is made of welded steel SP28 and consists of two half-shells, for the connection of which a wedge joint is used. In the place of the connector of the half-shells, a special assembly is mounted on which a charging charge is mounted, which is covered on the outer surface by special armor. On the flange of the front bottom, the final stage engine is mounted, operating through the main combustion chamber. On the rear of the engine there are four flanges for fixing the rocking nozzles, which provide control of the missile in the flight section of the I stage. As a power source of steering gears of swinging nozzles a solid fuel gas generator is used.
The rear section of the engine of the 1st stage is docked with a tail section, which has nodes for the longitudinal attachment of the missile to the TPK. As the nodes of the transverse attachment of the missile to the TPK, ring supports are used, dropped after the missile is thrown out of the TPK by means of a PAD mounted on the bottom of the TPK.
The TPK case, made of welded from shells and frames of aluminum alloy AMg-6, has brackets on the outer surface near the frames, necessary for fastening and rigging, lifting to the vertical position and launching the rocket. The front end of the container is closed with a lid, dropped before launching the rocket.
At the command "Start", operations that precede the launch of the rocket begin: checking the on-board systems, switching the missile to an onboard power supply, etc. After about 3 minutes, after the "Start" command, the elongated cumulative charge of the TPC cover is undermined, the powder engine of the lid release is launched and the latter is separated from the container. After separating the container connector block and breaking the bolts of the rocket attachment to the TPK, a powder pressure accumulator is installed in the container, and when a pressure of 6x10 5 N / m 2 is reached in the subattack volume the rocket starts moving. The flight control of the missile was attained by injecting spent turbine gas into the diverging section of the four sustainer nozzle.
The shape of the powder charge of the pressure accumulator is selected in such a way that the indicated pressure in the sub-rocket volume during the movement of the rocket in the container is maintained constant. At the moment of exiting the TPK, the rocket reaches a speed of 30 m / s. At an altitude of 10-20m above the container cut, the first stage of the solid-fuel rocket starts. At the same time, the support rings are separated and the rocket connector unit is divided. The engine of the first stage operates approximately 58 seconds. When the pressure in the chamber falls to 5x10 5 N / m 2the powder engine of the final stage is started, which works until the fuel burns out completely.
After the second stage engine starts, at the exit of which to the mode of 90% of the rated thrust, the stages of the rocket are separated. In the case of using the "light" head part on the 56c of the second stage engine, the head fairing is reset. When the required combination of rocket motion parameters (speed, coordinates, etc.), which provides the given range of fire, is achieved, the control system gives the command to turn off the engine. At the same time, the head part is separated.
Before the release of the missile from the TPK. if necessary, it may be possible to abort the start-up. There is also the possibility of an emergency detonation of a missile in flight.
At the first stage of the missile, four rotary nozzles of a solid fuel engine are used as controls. The nozzles are rotated by hydraulic steering machines. Powder pressure accumulator is used for gas production. The control of the second stage of the rocket at the angles of pitch and yaw is carried out by injecting gas into the supercritical part of the LPRE nozzle. The second stage was designed and manufactured in ampulized form. The second stage of the roll angle is controlled by two pairs of tangentially installed control nozzles. For the operation of control nozzles and injection is used gas, taken after the turbine of the turbo-pump unit of the propulsion system of the second stage (turbogas). The gas supply to the injection and control nozzles is effected by gas distributors, which are driven by electric motors.
The missile is controlled by means of six control channels:
- the roll angle stabilization channel;
- lateral stabilization channel;
- a normal speed control channel;
- a longitudinal speed control channel;
- the flight range control channel (the control channel for shutting down the engine of the second stage and separating the head part);
- channel separation control stages.
Each of the first four control channels is a closed automatic control system that operates on the principle of eliminating the deviation between the current value of the controlled parameter and its program value. Operation of the fifth and sixth channels is carried out by an open circuit, i.e. when the necessary conditions are given, commands are given for the separation of stages, switching off the second stage engine and separating the head part.
The rocket implements the so-called "hot" separation of the stages, in which the separation of the first stage occurs after the start of the second stage engine. At the end of the first-stage engine, the rocket gains about 27 km. It is unprofitable to make the separation of stages at such a low altitude, because because of the large aerodynamic forces acting on the rocket, considerable efforts would be required to set the stages to a safe distance. In this connection, the stages are separated after the rocket reaches a height of ~ 40 km. During the rise to this altitude, the controllability of the missile is provided by an auxiliary engine, the powder propulsion engine of the final stage of the thrust, which is started after the burning of fuel in the engine of the first stage.
The separation of the head part is made at the end of the active part of the trajectory during the aftereffect of the thrust of the engine of the second stage. First, three bursting bolts are triggered, by means of which the head part is fixed to the instrument compartment, and then the rocket part of the second stage is braked by the exhaust gas of the oxidizer tank through two anti-nozzles located on the front bottom of the tank. The anti-nozzles communicate with the atmosphere through two hatchways in the body of the instrument compartment. Opening of nozzles occurs as a result of triggering of elongated detonating charges, driven by electric detonators. The manhole covers of the instrument compartment are knocked out by plugs that emanate from the nozzles. After opening the nozzle, a pyrovalve is triggered, through which the boost gas flows in a direction perpendicular to the longitudinal axis of the rocket.
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