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My Army Redstone Missile Days

Page 12



Appendix A:
The Redstone Missile In Detail

7th Army

Battery A, 1st Missile Battalion, 333rd Artillery
40th Artillery Group (Redstone)
Bad Kreuznach, Germany

The following description of the Redstone missile is compiled from my personal experience; and,

Redstone Electronic Materiel Maintenance Course (REMMC)
Training Documents:

THE REDSTONE MISSILE SYSTEM, January 1959, publication 3262
THE REDSTONE MISSILE SYSTEM, August 1960, publication L619

Publication 3262 was issued to me when I became a REMMC student.
Publication L619 was provided to me by the Staff of Morris Swett Technical Library, Fort Sill, Oklahoma.

L 619

Chrysler Corporation Missile Division document "THIS IS REDSTONE".



September 1960.

FM 6-25

This document reflects the organization, structure, and tasks of the Redstone missile unit configured as a Field Artillery Missile Group employed through mid 1962, at which time Redstone missile groups were reorganized as Redstone missile battalions.

March 1962.

FM 6-25

This document reflects the organization, structure, and tasks of the Redstone missile unit configured as a Field Artillery Missile Battalion. In mid 1962 Redstone missile groups were reorganized as Redstone missile battalions, and remained as such until mid 1964, at which time all Redstone missile units were deactivated.

December 1960.

FM 6-35

February 1962.

FM 6-35

December 1960.

FM 6-36

October 1963.




John W. Bullard

Historical Monograph
Project Number: AMC 23 M

15 October 1965

AMC 23 M

(193 Pages)


February 14, 1955

Gnd. Equip.

This document reflects the initial design thinking in 1955 about the ground support equipment thought be needed for the Redstone missile. The actual equipment employed with the tactical missile in the field commencing in June 1958 turned out to be quite different. Perhaps the most notable example of this is the replacing of the 25 ton heavy duty crane with the portable lightweight erector servicer equipment to raise the missile to the vertical firing position.

I have included this document for its historical significance, particulary as the document was issued under a cover letter by Wernher von Braun.

The document was provided to me by the Staff of Morris Swett Technical Library, Fort Sill, Oklahoma.


Introduction and Overview

During its deployment Redstone was the Army's largest, and until the fielding of the Pershing missile, longest-range tactically operational field artillery guided missile. I should clarify this statement somewhat. In the mid to late 1950's the Army Ballistic Missile Agency (ABMA) under the Dr. Wernher von Braun civilian team at Redstone Arsenal, Alabama developed the 1,500 mile range Jupiter Intermediate Range Ballistic Missile (IRBM). However, under a subsequent Department of Defense directive assigning all land-based long range strategic missiles to the US Air Force, Jupiter was ordered to be turned over to and deployed by the Air Force. The DOD decision also limited the US Army to deploying tactical nuclear missiles with a maximum range of 200 statute miles (later increased to 400+ statute miles with Pershing). Thus, Redstone was the US Army's largest operationally deployed missile.

US Air Force Jupiter missile squadrons were based in Italy and Turkey from 1961 to 1963 as strategic weapons aimed at the heart of the Soviet Union. The Air Force was a reluctant recipient and operator of the Jupiter, in that in parallel to the development of the Jupiter by ABMA the Air Force had developed and deployed its own IRBM, the Thor, to Europe and based it in the United Kingdom. Eventually, all Jupiters were withdrawn from service by April 1963, in part because they were then classified as obsolete, and also as part of the deal created behind the scenes by John F. Kennedy with the Soviet Union to help end the October 1962 Cuban Missile Crisis, as a secret response to the Soviet Union withdrawing its nuclear missiles from Cuba.

Redstone had a rocket-type propulsion system using two liquid propellants, alcohol and liquid oxygen, which were pump fed into the motor by a steam-driven turbine. The guidance system was an inertial type, preset prior to firing, which automatically guided the missile to the target and corrected the trajectory as necessary to insure an adequate circular probable error (CPE).



The major characteristics and physical dimensions of the missile were as follows:

Length: 69' 4", Diameter: 70".
Loaded weight: 61,700 pounds, Empty weight: 16,500 pounds.
Maximum range: 200 statute miles.

Propellants -
Oxidizer: Liquid oxygen, 25,000 pounds;
Fuel: 75% ethyl alcohol + 25% water mixture, 19,000 pounds;
Steam source: Hydrogen peroxide, 790 pounds.

Thrust: 78,000 pounds for 96 to 117 seconds.

Guidance: Inertial.

Warhead: Nuclear, 7,900 pounds total nose section weight.
Option of a 3.75-Megaton (MT) "High Yield" Device, or 500-Kiloton (KT) "Low Yield" Device.
Proximity or Contact.

Mobility: 100%

Missile Description


The missile consisted of a thrust unit to provide propulsion, and a body to provide control for the entire missile and deliver the payload to the target. The body was comprised of the warhead section (initially known as the nose unit) housing the payload, and the aft unit housing the guidance and control instruments. The thrust unit was comprised of the center unit housing the fuel and oxidizer tanks; and, the tail unit housing the rocket engine and supporting 4 fixed stabilizers, 4 movable rudders, and 4 jet vanes which extended into the rocket exhaust to provide control of the missile until its speed was sufficient to cause the external rudders to be effective. The rocket engine, which produced 78,000 pounds of thrust, and a propellant pumping system operated by hydrogen peroxide, a chemical, which readily decomposes into steam was located inside the tail section. The oxidizer tank, located above the tail section, was loaded with 25,000 pounds of liquid oxygen (LOX). The fuel tank was located above the oxidizer tank and loaded with 19,000 pounds of a 75% ethyl alcohol-25% water solution.

Shell Exploded View
Shell Skeleton View
Warhead Section
Aft Unit
Center Unit
Tail Unit

The warhead section and the aft unit were joined by 8 nondestructive bolts, 4 of which were seated and installed through 4 ball and socket fittings used for mating alignment. The thrust and body units were joined with 6 ball and socket fittings using 6 bolts containing internal explosive charges. Shortly after shutoff of the propulsion system, the 6 explosive bolts were destroyed and the 2 missile units were pushed apart by 2 air-loaded pistons. To keep the missile weight as low as possible, only the body unit was made strong enough to withstand the forces created when the body re-entered the dense lower atmosphere as it descended to its target. Upon re-entry the pressure on the body reached 95 psi, and the temperature rose to 1,000 degrees F on the forward portion of the warhead section. The entire body was constructed with a skin of alloy steel riveted and welded to an underlying framework comprised of former rings, bulkheads and stringers. The warhead framework members were made of alloy steel while the aft unit framework used aluminum. A silicon rubber gasket was installed at the warhead/aft unit mating interface to create an airtight seal for the pressurized instrument compartment.

Aft Unit Skirt Section
Aft Unit
Center Unit
Center Unit

The lower portion of the body, the aft unit, housed the guidance system. A reinforced pressure bulkhead divided the aft unit into two sections: the instrument compartment and the skirt section. The guidance system, located in the instrument compartment, was an inertial type, which operated on information from accelerometers, which measured the movements of the missile. The accelerometers were mounted on what I would call the heart and soul of the guidance system; a three degrees of freedom stabilized platform, the ST-80. The ST-80 was kept in proper orientation by three gyroscopes. The relative movement between the missile and the ST-80 measured the angular position of the missile about its center of gravity (cg). An onboard tape recorder contained most of the trajectory data, although some data was set directly into the computers. Four vanes located at the base of the body controlled the body during reentry phase into the lower thicker atmosphere.

Tail Unit
Tail Unit
Rocket Engine

The warhead section, weighing 7,900 pounds, contained the nuclear warhead. In print and on the Internet there are various versions of what warhead was carried by the Redstone missile. This is what we were taught and told at the time of my service with Redstone. There were two warhead options for the Redstone missile that could be installed inside the nose unit, a so-called "high yield" 3.75-megaton thermonuclear device, and a so-called "low yield" 500-kiloton thermonuclear device.

Since the choice of warhead yield had nothing to do with the external shape and dimensions of the nose unit, the missile could be equipped with one or the other, depending on mission objectives. Choosing which one to carry was certainly not a function or decision of the Firing Battery. During missile vertical checkout, using the warhead-arming panel in the FC&TT, the warhead could be selected for either an air burst (proximity) or a ground burst (impact). At the time the air burst was commonly referred to as "clean", and the ground burst as "dirty".

The Redstone Propulsion System.


Propulsion 2
Propulsion 1
The Redstone utilized a bi-propellant, pump-fed, liquid rocket propulsion system. The components were located in the thrust unit. Redstone used an alcohol-water solution for fuel with liquid oxygen (LOX) as the oxidizer. A steam driven turbine powered the propellant pumps. The steam was created by the decomposition of the hydrogen peroxide into super-heated steam, which was routed through turbine blades to drive the propellant pumps. Various propulsion system control valves were operated by high-pressure air. High-pressure air was also used to maintain a slight positive pressure in the fuel tank. The oxidizer tank was self-pressurized by the evaporation of LOX. Initial ignition was accomplished by injecting igniter alcohol from an external source into the rocket engine combustion chamber, where it was combined with LOX. An electrically energized igniter squib ignited the mixture. With successful initial ignition, alcohol from the missile fuel tank was fed to the combustion chamber.

The design of the combustion chamber and nozzle of the rocket motor was such that thermal energy of the burning propellants was converted to kinetic energy. The thrust, or forward propulsive force, was proportional to the kinetic energy. The thrust produced was held at a constant value by maintaining a constant pressure within the combustion chamber. Any change in the combustion chamber pressure was sensed and converted into an electrical signal that changed the speed of the propellant pumps. In this manner the quantity of propellants consumed each second was changed to the value which created the correct chamber pressure and thrust. Thrust was terminated at a preprogrammed time by stopping the flow of propellants to the rocket motor and stopping the flow of hydrogen peroxide to the steam generator.

The Redstone propulsion system was comprised of the following major components and systems: pneumatic system, electrical system, rocket engine, fuel system, oxidizer system, steam generating system, and ignition system.

Propulsion Pneumatic System

A supply of high-pressure air was carried in six spheres mounted in the tail section. Prior to liftoff the spheres were charged to 3,000 pounds per square inch (psi). Air from these spheres, after passing through the heat exchanger maintained slight pressure in the alcohol tank. High-pressure air was also supplied to a pneumatic panel, controlled by regulators and valves, and distributed to pressurize the hydrogen peroxide system and to operate various propellant controls valves.

Propulsion Electrical System

The electrical power required by the propulsion system was derived from three sources:

Batteries: batteries located in the missile body guidance compartment supplied 28-volt direct-current (d-c) power. The power was routed to various relays and rocket engine controls, which provided automatic sequencing of engine ignition and shutoff operations.

Ground source: 120-volt, 60 Hz alternating current (a-c) power was supplied by a diesel generator ground source prior to firing, to operate various heaters in the propulsion system.

Inverter: 115-volt, 400 Hz alternating current (a-c) was supplied by an Inverter in the missile body unit to operate the thrust controller. The inverter transformed the 28-volt d-c from one of the missile batteries to a-c.

Rocket Engine

The Redstone missile was powered by a North American Aviation Company Rocketdyne Division NAA 75-110 A-6 (and later the A-7 final version) liquid bi-propellant rocket engine rated to provide a nominal thrust of 78,000 pounds. The capacity of the missile propellant tanks limited burning time to a maximum of 117 seconds. Operation was by continuous injection and combustion of fuel and oxidizer. The rocket engine was very simple in construction. It had a large, cylindrical, double-walled combustion chamber, open at one end for escape of powerfully expanding gases, and closed off at the forward end by a perforated injector plate. Alcohol and LOX were forced under pressure through the perforations, and atomized, mixed, and ignited just below the plate. Started electrically, ignition occurred along a broad flame front covering the full cross-sectional area of the chamber. The burning gases expanded violently and gained velocity, rushing to escape through the narrow throat of the chamber. Additional thrust was delivered when the gases surged into the flaring exhaust nozzle, expanded still more, and emerged as a white-hot jet stream.

Fuel System

Thrust Unit
The alcohol tank was filled with 19,000 pounds of a 75% alcohol-25% water solution when the missile was in the vertical position. The 25 percent water content of the fuel reduced the flame temperature so that the engine would not melt; and, added to the weight and pressure of the gases expelled, thus contributing to thrust. For conditions where it was known that the ambient outside temperature would be above 35 degrees F at time of launch, prior to alcohol loading, 10 gallons of water used as an inert lead start were placed in the rocket motor manifold. For conditions at or below 35 degrees F lithium chloride was employed. Lithium chloride had a freezing point of -105 degrees F. The lithium chloride was transported in a 20-gallon tank contained in the alcohol trailer. A motor-driven pump with hose attachments was used to feed the lithium chloride into the regenerative jacket and manifold of the engine.

During ignition, the water, or the lithium chloride, was forced into the combustion chamber ahead of the main alcohol flow to reduce the violence of main-stage ignition. The alcohol tank was pressurized to 20 psi by air from the high-pressure spheres. The turbo-pump drew 150 pounds of fuel per second from the tank and forced it through dual ducts to a manifold encircling the engine exhaust nozzle. The alcohol then flowed upward between the engine double walls, simultaneously cooling the engine effectively and raising the alcohol temperature to make it easier to ignite. The alcohol finally passed through the injector plate and into the combustion chamber.

Oxidizer System

With the missile in the vertical position, the LOX tank was filled with 25,000 pounds of LOX. Because of its extremely low boiling point temperature of minus 297 degrees F, the LOX evaporated rapidly. Due to the high evaporation rate, LOX tank replenishment was required just before the missile was fired. LOX was drawn from its tank by a turbo-pump and delivered to a small reservoir known as a "LOX dome" above the engine injector plate. The normal flow rate was 200 pounds per second. The LOX tank was pressurized to 30 psi to ensure smooth flow through the turbo-pump. During engine starting sequence, this pressure was provided by a ground source of compressed air. When the engine began to operate, the 30-psi pressure level was continuously maintained by gaseous oxygen. For this purpose, a small amount of LOX was bled from the main LOX line and passed through coils in the exhaust duct of the steam system, where heat converted it to gaseous oxygen.

Steam Generating System

Hydrogen Peroxide (H2O2) produced steam to power the turbo-pumps that delivered propellants to the engine. A 72-gallon hydrogen peroxide tank located just forward of the engine was filled when the missile was vertical and was pressurized by the pneumatic system. When a feed valve was opened, hydrogen peroxide flowed to a steam generator containing potassium permanganate pellets. Upon hitting the bed of pellets, the hydrogen peroxide, undergoing a process known as a catalytic conversion, was instantaneously converted into super-heated steam. The steam rotated the turbo-pumps and exited through an exhaust duct containing heat exchanger coils. Some of the heat was recovered here to maintain the pressure level in the pneumatic system and the LOX tank.

Hydrogen Peroxide
Steam Generator

The escaping steam contributed slightly to missile thrust. Thrust was maintained at a constant value by a servomechanism known as the thrust controller. The thrust controller measured pressure in the combustion chamber, compared it with a desired target value, and acted, according to the difference, to either increase or decrease the flow of hydrogen peroxide to the steam generator. In this manner, steam production was adjusted to alter the speed of the propellant pumps, and thereby control the rate of propellant flow to the rocket engine.

Ignition System

Once started, combustion was self-sustaining. However, because alcohol and liquid oxygen do not ignite spontaneously when mixed, a method of initiating combustion had to be employed in the rocket engine. This ignition system consisted of the following:

Igniter Alcohol: an igniter alcohol supply, consisting of a 3.5-quart container mounted on the missile launcher, was pressurized during the missile preparation sequence. Alcohol from this container was forced at the proper time through the valve box and multiple coupling to the center ring of the rocket engine injector plate, which functions much like a showerhead.

Igniter Cartridge: an igniter cartridge assembly, consisting of two electrically fired pyrotechnic squibs with a burning time of 10 seconds, initiated combustion. This assembly was suspended beneath the injector head by means of a thin plastic rod that was screwed into the injector head prior to firing.

Ignition Sensing Device: An ignition sensing device, commonly called the mainstage stick, was installed below the rocket engine nozzle. The mainstage stick had a loop of wire, which extended into the jet stream of the engine. With proper ignition, the wire burned in two, generating an electrical signal that caused the main alcohol valves to open and the turbo-pump to activate.

Rocket Engine Operation

The following events occurred automatically upon actuation of the firing circuit from the remote-firing panel. Some of these events occurred simultaneously.

  • Alcohol and hydrogen peroxide tanks were pressurized. However, from the time the remote-firing panel fire switch was activated, it took approximately 7.5 seconds for the LOX tank to pressurize.

  • With alcohol tank pressurization, LOX tank pressurization from the ground high-pressure air supply began.

  • The igniter squibs were energized, and the cartridge began to burn in the igniter assembly. The igniter link burned through.

  • The main LOX valve opened. LOX flow to the rocket engine was caused by a combination of gravity and tank pressurization.

  • Alcohol from the igniter alcohol container was forced through the center ring of the injector head of the combustion chamber.

  • In the presence of excess oxygen, combustion began.

  • The mainstage stick sensed ignition and caused the main alcohol valves to open. Alcohol then flowed into the manifold and up through the combustion chamber walls, forcing the inert lead start into the chamber.

  • The hydrogen peroxide valve opened to admit this chemical into the steam generator.

  • Super-heated steam was produced to drive the turbines, and the turbine pump was accelerated towards its rated speed.

  • Thrust rose rapidly as the flow of alcohol and liquid oxygen increased.

  • The admitting of gaseous oxygen into the LOX tank maintained constant pressurization.

  • The turbo-pump attained its rated speed of 4,800 rpm.

  • Full thrust was developed. When the thrust developed exceeded the weight of the missile, the missile left the launcher.


At the correct time predetermined by the range to the target, thrust was terminated by stopping the power to the turbine, and therefore stopping the flow of propellants to the engine. After motor cutoff, the propulsion system served no further function.

The Redstone Guidance System


Guidance 2
Guidance 1
The Redstone guidance system maintained proper angular orientation of the missile; sensed, measured, and corrected deviations from a predetermined flight path; and, determined the point at which thrust was terminated. The guidance system was of the inertial type; that is, a system that was completely self-contained, and depended on no other information other than errors generated as a result of positive or negative changes in velocity. These velocity changes occurred in either one of the two measuring planes: a lateral plane or a range plane. The missile guidance system had two types of control: attitude control and path control.

Attitude Control: Attitude is the angular position of a vehicle with respect to its center of gravity (cg). The terms used to define attitude are roll, pitch, and yaw. Roll is the angular rotational motion and orientation of the missile about an axis drawn longitudinally through the center of the missile from its nose to its tail. Roll can be clockwise or counter-clockwise as viewed along the roll axis. With the missile in a horizontal position, yaw is the motion and orientation of the missile about an axis perpendicular to the horizon and through the cg of the missile. Yaw may be clockwise (to the right) or counter-clockwise (to the left) of the desired orientation, as viewed from above the missile. Pitch is the motion and orientation as seen from the side about the pitch axis through the missile cg, and may be up or down.

Path Control: Guidance or path control is the determination and control of the missile's position with respect to its references. Range guidance is in the direction of the target, and in Redstone, was measured along a line perpendicular to the tangent of the trajectory at impact. The other reference direction is lateral: the displacement right or left of the azimuth between the launcher and target.

Components: To satisfy the requirements for attitude and path control, self-contained equipment measured the performance of the missile, determined the amount of deviation from the desired conditions, formed corrective commands, and repositioned the missile as necessary. The Redstone guidance system consisted of a gyroscope stabilized platform, the ST-80, as a reference, accelerometers that measured performance, computers that determined corrective commands, a relay box that applied battery power to motor actuators. The motor actuators positioned the rudders and steering vanes as required. Motor actuator feedback circuits provided stability and prevented over-control. The heart of the system was the ST-80. It was automatically leveled and aligned to a fixed reference position before firing. Three gyroscopes mounted on the platform maintained the alignment throughout the flight all the way to impact.

Attitude Control

General: One of the guidance requirements, that of attitude control, was accomplished by potentiometers (voltage measuring devices) mounted between the ST-80 and the Missile frame. If the missile developed a roll, pitch, or yaw error, the angular error between the ST-80 and the missile frame was electrically measured by the potentiometers, and their electrical output signals were fed into the control computer. The error signals were mixed in the control computer to produce voltage commands, which were used to reposition the missile. During powered-flight phase, the combined effects of the jet vanes in the exhaust stream of the rocket motor and air rudders on the thrust unit produced the necessary control torque to reposition the missile. During travel through the midcourse portion of the trajectory, which is essentially out of the atmosphere, a system of air jet nozzles using high-pressure air produced the necessary control forces. During the terminal portion of the trajectory, air vanes located on the aft end of the body produced the attitude control.

Pitch Programming: Redstone, launched from a vertical position, was steered into a ballistic trajectory. A timed pitch program was introduced into the control system to cause the missile to assume the correct attitude, which was approximately tangent to the trajectory. Timing signals to effect attitude changes and other functions along the missile trajectory were stored in the form of pulses on a magnetic tape. The tape was run through a tape playback unit in the guidance compartment. The pulses on the tape were sequentially fed into the ST-80 to initiate the action of a stepmotor. With each pulse, the stepmotor displaced the wiper arm of the ST-80 pitch potentiometer from its null position in a direction opposite and away from the direction to the target. This action was sensed as an error by the ST-80. To correct the error and bring the wiper back to its null position, the missile was correspondingly pitched over about its cg towards the target. Step by step the missile was literally commanded to rotate nearly 180 degrees about the ST-80 in its fixed reference position. At impact the missile was in essence upside down while the ST-80 was still right side up.

Since the program was different for various trajectories, the proper program corresponding to a desired trajectory was imposed on the missile tape recorder from a program recording system in the FC&TT. Tapes for different trajectories were from a library of tapes carried by the Firing Battery, residing in a safe inside the FC&TT. The stored trajectory tapes were Mylar punched tapes, which were fed through a punched tape reader mounted in a FC&TT console. The firing data pulses read off the Mylar tape were recorded on the missile magnetic tape recorder.

Jet Nozzle System: Because the Redstone passed through areas in which there is little atmospheric air, there was essentially no control exercised through the use of moveable control surfaces. Consequently, attitude control had to be accomplished in another manner. A jet nozzle system was employed. Two jet nozzles mounted opposite to each other were located at the base of each of the four body unit air vanes. When the air vanes moved to correct an attitude error, the respective air jet nozzles were activated, creating a small thrust by exhausting high-pressure air, rotating the missile about its center of gravity to the proper attitude. For example, if a pitch up error existed, air jets on the opposite side of the missile were activated, exhausting air downward to create an upward reaction, which rotated the missile to the proper attitude.

Path Control

The inertial system sensed errors from a predetermined reference trajectory. The path of the missile was corrected during the latter portion of flight so that the actual impact point would coincide with that of the reference trajectory. The adjustments were accomplished by signals generated by two accelerometers mounted on the ST-80. These accelerometers were air bearing gyroscopic devices with pendulum masses and were single integrating mechanisms, which sensed accelerations and converted them to velocity outputs, which were fed to the range and lateral computers. The source of the air was the air pressurized to 3,000 psi carried in one of the two air spheres located in the skirt section of the aft unit.

The range and lateral computers were unique devices, actually electro-mechanical analog computers. To attain displacement information, the velocity signals were again integrated to displacements, this time by ball and disc integrators in each computer. One way to describe a ball and disc integrator is to liken it to a phonograph turntable, the disc being the phonograph record, and the ball being the needle and phonograph cartridge at the end of the phonograph tracking arm. The displacement control signals from the range and lateral computers were then fed into the control computer. The control computer then sent correction commands to the appropriate set of steering vanes, and associated air jet nozzles when the body unit was still above effective atmosphere. The feedback system prevented over-control, or over steering the body unit.

Range Control: The range accelerometer was oriented on the ST-80 so that it measured in a direction perpendicular to the trajectory tangent, e.g., the impact coordinate, and in the plane of the trajectory. In this way, it gave velocity information pertaining to the location of the missile compared to the reference trajectory. Its output was fed to the range computer which performed the second integration for obtaining displacement, computed a corrected thrust termination time, and supplied arming signals to the warhead fuze. Any range errors that existed at cutoff, as well as deviations caused by various disturbances after cutoff, were measured, stored, and sent out as path corrections during the terminal guidance phase.

Lateral Control: The lateral accelerometer measured in the direction perpendicular to the plane of the trajectory and gave information as to the location of the missile compared to the plane of the reference trajectory. Deviations were measured along the entire trajectory and corrective signals were introduced into the control system of the missile during the initial and terminal phases of the flight.


Since all path control information was determined by devices wholly within the missile, a precise degree of alignment of the ST-80 on the firing azimuth was necessary. Fin I of the missile was always pointed toward the target. Two Theodolites and a missile prism were used to establish the precise firing azimuth by referencing to a previously surveyed orienting line. The missile prism was a reflector mounted on the missile and was aligned to the missile airframe, as was the ST-80.

The firing angle was established, and the missile then in the vertical position, was rotated until the missile prism was sighted from one of the Theodolites. After final adjustments, such as leveling the launcher within a close tolerance, final laying of the missile was carried out. Since the missile prism was mechanically aligned to the ST-80 and since the missile had been rotated as a result of sighting on the prism from the Theodolites in reference to the firing azimuth, the range and lateral accelerometers of the ST-80 were also oriented with respect to the target.


During the preliminary design of the missile, it became apparent that it would be advantageous if the body and thrust unit were separated after thrust termination. The high dynamic pressure that the missile would encounter during descent toward the target became an important consideration in the design of the missile from structural, aerodynamic, and controls considerations, especially if the missile maneuvers during terminal guidance. A large amount of structural weight in the thrust unit could be saved if it were constructed to meet only the conditions encountered during the ascent phase of flight when missile speed is relatively slow.

A separate re-entry section offered a more favorable center of pressure location over a wide range of velocities encountered, thus improving stability. Control forces were less for the short lighter body. Therefore, the missile design incorporated the separation of the body from the thrust unit after the powered flight phase was completed. The two units were connected until cutoff by six bolts, each containing an explosive charge in the bolt head. Shortly after cutoff, at the point where forward acceleration of the missile decayed to zero and the missile now travelling at constant velocity, the program computer supplied a signal, which detonated the six bolts. At the same time two expulsion cylinders, located at the top of the thrust unit and positioned 180 degrees opposite each other, were activated to supply the force necessary to push the two units apart. The expulsive cylinders were powered by high-pressure compressed air. The two units were separated at zero acceleration to prevent the now-separated thrust unit from accelerating into the body unit.

Temperature Control

Three temperature maintenance systems were used in the missile. The entire instrument compartment was heated or cooled as required to maintain a compartment temperature of about 60 degrees F prior to firing the missile. This was accomplished thought use of an external "drop tank" attached to the outside of the instrument compartment. During my time with Redstone the drop tank used contained a heater and a blower, and initially was packed with dry ice.

The drop tank was eventually upgraded to one that used liquid nitrogen rather than dry ice. I believe the changeover was incorporated in the first half of 1961. Liquid nitrogen was a by-product of the manufacture of liquid oxygen by the Engineer Company.

An additional cooling system, into which cubes of dry ice were packed, was located inside the instrument compartment to maintain a uniform ambient temperature in the jacket surrounding the ST-80 stabilized platform during the trajectory. The ST-80 contained heating elements that helped preheat it. The heat generated by the ST-80 in operation was sufficient to keep it at its required operating temperature.

Power Supplies

The Redstone required four types of electrical power: for its operation.

28-volt d-c: Power to operate the rudder actuators, various relays, valves, ignition circuits, and the 400-Hz a-c inverter was obtained from two high capacity 28-volt d-c nickel cadmium (nicad) batteries.

60-volt d-c: Power used in the guidance measuring and feedback circuits was provided by one 60-volt d-c battery.

115-volt 400-Hz a-c: Power for amplifier circuits and in the ST-80 stabilized platform was provided by the 115-volt 400-Hz inverter powered by one 28-volt d-c battery.

208-volt and 115-volt 60-Hz a-c: Power to operate the d-c generators for use in testing and firing the missile, and to operate missile heaters prior to launch, was provided by a 60-kilowatt, 208-volt 60-Hz trailer-mounted diesel generator. This generator also supplied power to operate propellant transfer equipment.

Trajectory Sequence

The Redstone missile followed a modified ballistic path, divided into four phases.

Phase I: The first phase of flight was the powered phase during which the rocket engine operated. During the first few seconds after liftoff, 4 ablative carbon jet steering vanes extending into the jet exhaust accomplished control of the missile's flight path. Once sufficient velocity, and therefore sufficient air flow over the four moveable steering rudders at the base of the thrust unit was attained, the rudders in conjunction with the jet vanes steered the missile. The mission pitch program inserted into the guidance system began to pitch the missile toward its target. Attitude control and lateral control were exercised and maintained. Range information was used to determine the precise engine cutoff point. When missile velocity and displacement for the missile to coast to its target were reached, the guidance system initiated action to shut down the propulsion system.

Phase II: The second phase of flight was the period between engine cutoff and missile separation, during which the missile settled to a free-flight condition. A signal from the control computer initiated separation, by exploding the six bolts holding the two units together, and by pushing the two units apart by the air-loaded pistons.

Phase III: The third phase of flight was the period between separation and reentry. The expended thrust unit followed its own ballistic path, albeit shorter and uncontrolled, to eventual impact destruction about 10 statute miles short of the target. The body continued to coast to its target. Body unit attitude control was maintained, but range and lateral errors were determined and stored for correction just before impact The missile was essentially above the earth's effective atmosphere during this phase, travelling at high supersonic speed, and reaching a maximum altitude of approximately 60 statute miles. Since body steering vanes were ineffective in this rarified atmosphere, the air jet nozzles were employed to maintain attitude control. Two or more were activated as required, generating thrust in opposition to missile movement, in order to re-orient the missile along its flight path.

Phase IV: The fourth and final phase of the trajectory began at what is called re-entry, the point about 20 statute miles altitude where the missile encountered enough air resistance to cause it to slow down, although it was still travelling at supersonic speed. The denser atmosphere activated a pressure switch, which then allowed for the stored range and lateral errors to be corrected.

The Redstone Handling Equipment

Missile Transporters

The packaging and transportation equipment permitted the missile and ground equipment to be moved by road, rail, water, and air. The missile was shipped from the Chrysler Company plant to 40th Artillery Group in the same container and was checked out by the 630th Ordnance Company without having to be completely removed from the container. The ST-80 stabilized platform was packaged separately and was shipped from the Ford Instrument Company to the Ordnance Company. After conducting acceptance tests, the Ordnance Company forwarded the ST-80 to the Firing Battery for installation in the missile. A warhead shelter was carried in a winterization vehicle. During inclement weather, the shelter was used to protect the ST-80 during installation, and also when work was performed at the guidance or warhead compartments.

Missile transporters included:

Warhead Unit Semi-trailer: This was a single-axle 2-wheel vehicle to provide storage protection and transportation for the warhead unit, van cover jacks, and all explosive items and devices used in a launch operation.

Aft Unit Trailer: The aft unit trailer was a single-axle 2-wheel special semi-trailer that provided storage protection and transportation for the aft unit.

Thrust Unit Semi-trailer: This single-axle 2-wheel special semi-trailer provided storage protection and transportation for the thrust unit, and storage and transport space for the van cover jacks, and missile components which were installed at the firing site.

Thrust Unit
Thrust Unit
Thrust Unit

Stabilized Platform and Carbon Dioxide Truck

This dual purpose 2&1/2-ton 6x6, M35 truck initially provided transportation for 2,000 pounds of solid carbon dioxide (dry ice) in skid -mounted storage containers. Dry ice was originally used for temperature control of the guidance system compartment. The ST-80 was also carried on the truck in a controlled container, since d-c power was available from the vehicle to heat the container as necessary.

In 40th Artillery Group, 580th Engineer Company produced the dry ice initially needed for a Redstone launch operation. The dry ice was shipped to the Firing Batteries in 1-ton containers. It was then sawed into 1-inch cubes and packaged in 50-pound insulated bags that could be handled by one man. The 1-ton container was used to store the bags until they were needed. The dry ice was transported to the launch site on the ST-80 stabilized platform and carbon dioxide truck. At the launch site the cubes of dry ice were placed in the ST-80 inner cooler and in the external drop tank. Both tanks were filled to capacity, with the ST-80 cooler holding about 18 pounds, and the drop tank about 150 pounds.

Liquid Nitrogen

Nitrogen Trailer
Liquid nitrogen was also a by-product of liquid oxygen production by 580th Engineer Company. As a consequence, sometime in late 1960 or early 1961, the drop tank system was upgraded to employ the use of liquid nitrogen in lieu of dry ice. The liquid nitrogen was stored and transported in special tanks carried on a 2-wheel trailer. The trailer was towed behind the 3/4-ton Hydrogen Peroxide truck. A flexible metal hose was connected to the drop tank from the liquid nitrogen trailer.

Checkout and Firing Equipment

The checkout and firing equipment was designed to carry out functional tests and inspections of the missile to the extent necessary to insure the success of the mission. Firing Battery troubleshooting was limited to tasks that indicated defective items that could be easily replaced in the field. Typical were amplifier boxes, solenoid valves the range and lateral computers, and the ST-80. These and certain other replaceable missile and ground equipment components were carried in the Battery's spare parts truck.

Checkout and firing equipment included:

Fire Control and Test Truck (FC&TT)

The Fire Control and Test Truck, a 2&1/2-ton, 6x6, shop van, accommodated a number of equipment racks and test panels for prefiring tests of the missile during a firing operation. It also was used for performing periodic checks of missile components and systems maintained on a standby basis. The FC&TT also served as the center of the firing site communications. Additionally, the FC&TT was sometimes used by the Ordnance Company to carry out more detailed and thorough checkout and troubleshooting procedures.

Battery Servicing Shop

The battery servicing shop was a 3/4-ton, 2-wheel trailer used for storing and transporting missile batteries, battery service and test equipment, and other accessories to the launch site. Battery activation tasks were performed within the trailer.

Generator Trailer

The 2&1/2-ton, 2-wheel generator trailer provided the portable base for mounting the Cummins diesel generator and accessory equipment. The generator provided 120-volt single-phase and 208-volt 3-phase, 60-Hz electrical power to the power distribution station.

Power Distribution Station

The power distribution station was a 3/4-ton, 2-wheel trailer to convert 208-volt, 3-phase, 60-Hz alternating current from the generator to 28-volt direct-current, 60-volt direct-current, and 400-Hz electrical power used in the missile and other ground equipment during missile checkout.

Accessories Transport Truck

The 2&1/2-ton 6x6 accessories transport truck served as the prime storage and transport vehicle for loose items of equipment, and accessories necessary for missile checkout and servicing. Power and signal cables reels were carried on this vehicle.

Air Compressor Truck and Air Servicer Trailer

The 2&1/2-ton, 6x6, truck served as a base for mounting the air compressor, regulating system, and accessory equipment. The compressor provided compressed air, nominally at 3,000 psi maximum, at a specified dew point, to the missile for testing and for pressurizing the onboard storage spheres before firing.

Initially, high pressure air was taken directly from a 1.5 cubic foot air storage bottle mounted on the truck. After Redstone was deployed, this system was upgraded to employ an air servicer trailer. The trailer contained high pressure air cylinders which were pressurized by the compressor truck. The output of the service trailer air cylinders were connected to the missile, and remained so until launch.

Fire Truck

A 2&1/2-ton fire truck, towing a water tank trailer, was employed for emergency fire fighting, and for diluting or flushing away propellant spillage. The truck had an internal water tank storage capacity of 1,000 gallons and a pumping capacity of 1,500 gallons per minute.

Water Tank Trailer

This 4-wheel trailer furnished 2,000 gallons of reserve water to the fire truck. The truck and trailer supply permitted pumping at maximum capacity for 2 minutes.


The 2&1/2-ton 6x6 erector-servicer truck housed a 10-ton winch and 1-ton electric hoist. The truck carried the erector-servicer equipment consisting of: the fixed H-frame and H-frame spreader bars; the rotating A-frame; jack-type support stands; a hydraulic cart with arresting cylinders; a winch; and, erecting cables with associated pulleys and blocks, all compactly stowed on the truck bed. The purpose of the H-frame was to prevent the movement of the truck with respect to the launcher during assembly and raising of the missile. The H-frame was also employed as a boom for supporting and positioning a service platform used to access the missile body compartments with the missile in its vertical firing position.

The A-frame was employed as a boom for lifting and positioning the aft unit to the warhead unit, for lifting and positioning the rotating frame assembly to the thrust unit, and for lifting and suspending the thrust unit for mating to the body. The rotating frame assembly, commonly called the "tilt ring", was secured to the rear of the thrust unit and attached to the launcher. The tilt ring served as a hinge for raising the missile.

When the missile was ready to be raised to its vertical position, the A-frame, with cables attached, was approximately in the vertical position. Winch power was then applied to the A-frame, which pulled the missile to the vertical by functioning as a lever arm for the force transmitted though the cables. As the A-frame pivoted downward on the launcher in the direction of the erector truck, the missile was raised to the near-vertical position. The two hydraulically operated pistons attached to the launcher engaged the tilt ring, and slowly lowered the missile to its fully vertical position. After the missile was vertical, rollers on the tilt ring permitted the missile on the launcher to be rotated to the correct firing azimuth.


The launcher was towed to the firing site behind the erector-servicer truck. The launcher consisted of a base, an exhaust flame deflector plate, and the "tilt ring" rotating frame assembly. The launcher was mounted to a removeable single axis at its outrigger support arms, which provided a means of quickly disconnecting the wheel and axle assembly once the launcher was emplaced at the firing position.

Propellant Vehicles

Propellant handling and transportation equipment was designed and selected for long distance road travel and limited cross-country movement. The transfer equipment filled the missile tanks rapidly and conveniently. The regular electrical power source at the firing position was used to drive the propellant pumps. The liquid oxygen was pumped from the generating plant where it was produced directly into the trailers that delivered it to the firing position. Hydrogen peroxide was carried in the original container supplied by the manufacturer. Alcohol was normally shipped to the Ordnance Company in 55-gallon drums, where it was transferred into a trailer and mixed with water.

The 2-wheel, 9-ton special tank, liquid oxygen semi-trailer provided storage, transport, and pumping ability for liquid oxygen. The LOX transfer equipment was mounted in a closed compartment at the rear of the trailer. The LOX semi-trailer provided initial filling of the missile tank, and final filling to replenish evaporation losses.

The 2-wheel, 3,000 gallon capacity alcohol tank semi-trailer, resembling a gasoline tanker of that era, provided storage, transport, and pumping ability for the 75-percent alcohol, 25-percent water fuel mixture. The fuel transfer equipment was mounted in a closed compartment at the rear of the trailer, and provided the means of metering, filtering, and transferring the fuel to the missile tank.

The 3/4-ton, 4x4, hydrogen peroxide servicer truck provided transport and the means of handling a 78-gallon drum of concentrated hydrogen peroxide. The vehicle maintained the hydrogen peroxide at a temperature of 75 degrees F, +/- 10 degrees F, by using heater and cooler kits provided with the truck. Pumping equipment connected to an a-c power source was used to transfer the hydrogen peroxide to the missile at the firing position.

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