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Reentry Vehicle Development Leading to the Minuteman Avco Mark 5 and 11.

The Minuteman Intercontinental Ballistic Missile (ICBM) has been deployed for over fifty years. As one of two second generation ICBMs, Minuteman represented the ultimate solution to the concept of land-based offensive strategic weapons. The solid propellant propulsion system provided for a nearly instantaneous response while reducing maintenance efforts and costs significantly below those of the first generation cryogenic oxidizer Atlas and Titan I. Even the second generation Titan II with its storable liquid propellants and comparable response time was cumbersome in comparison.

Development of the business end of all the ICBMs, the reentry vehicles, likewise went from the first generation heatsink thermal protection system to the second generation ablative reentry vehicles enabling larger payloads (the reentry vehicle was lighter) to be carried as well as improving accuracy. This article discusses the evolution of reentry vehicle design and fabrication leading up to and including the Minuteman Mark 5 and Mark 11 reentry vehicles. Detailing the earliest efforts of the Army, Navy and Air Force reentry vehicle approaches puts the development of the Minuteman Mark 5 and Mark 11 reentry vehicles into the proper historical perspective. The discussion of the Army's effort covers only the Jupiter IRBM program and its pioneering work on ablation. The Navy's contribution was a much different approach to the heatsink concept with the discussion ending with the Polaris A-l and A-2 as the follow-on programs closely resembled the later Air Force approach. Due to classification issues caused by current world events, the third generation Air Force reentry vehicle designs are not discussed in this article though they have been described in great detail in an earlier article by Lin. (1)

Early Research

While bombardment rockets have been used for centuries, it was not until the creation of the German V-2 (also known as the A-4) that the warhead needed thermal protection due to reentry into the Earth's atmosphere. (2.) Since the entire V-2 impacted the target, there was no true separable reentry vehicle. (3) The original design called for the use of a lightweight alloy of magnesium and aluminum but wind tunnel tests indicated that from an altitude of 43 nautical miles, the operational maximum altitude, reentry into the lower atmosphere at 3,345 miles per hour would result in a warhead compartment skin temperature of 1,250 degrees Fahrenheit. Therefore the decision was made to use 1/4 inch sheet steel resulting in the need to decrease the explosive payload to hold the total warhead weight to 2,200 pounds (the steel casing weighing 550 pounds). The explosive chosen for the warhead was Amatol, a mixture of sixty percent TNT and forty percent ammonium nitrate, which was insensitive to heat and shock. There was no warhead compartment insulation. (4)

Arming a guided missile derived from the V-2 with an atomic warhead was an obvious next step in strategic warfare since it was only a matter of time for atomic bomb design to catch up with guided missile delivery capability. Concerned with the vulnerability of the eastern United States to long range missiles from the Soviet Union, in 1945 the National Advisory Committee for Aeronautics (NACA) realized an urgent need to begin studying the problems of hypersonic flight (defined as greater than five times the speed of sound which is the speed at which aerodynamic heating begins to be significant). By the late 1940s, two major NACA facilities, Ames Aeronautical Laboratory (Ames), Moffett Field, California, and Langley Aeronautical Laboratory (Langley), Hampton, Virginia, responded by expanding their aeronautical work to study aerodynamic issues involved in ballistic missile flight. (5)

Theoretical research into the problem of aerodynamic heating of ballistic missiles upon reentry into the atmosphere at high speeds was first published in 1949 by Carl Wagner. (6) The first comprehensive theoretical work was begun in 1951 by H. Julian Allen and A.J. Eggers, Jr., engineers at Ames. They studied the problem of reentry heating for ballistic, glide and skip-entry trajectories. Their investigation of the three types of trajectories was driven by the need to find a flight path that could best utilize the thermal protection materials then available. Allen and Eggers dismissed the pointed nose shape, a carry over from rifle bullet design, at the start, instead focusing their calculations on a blunt, hemispherical shape, recommending that "not only should pointed bodies be avoided, but that the rounded nose should have the largest radius possible." (Figure 1)

It is important to note that these calculations were made with "light" and "heavy" missile options and no mention was made of a reentry vehicle as such. The "light" missile optimum nose shape from a heat transfer standpoint was a blunt shape; for the "heavy" missile a more slender shape was optimum. Their calculations showed that the high drag caused a detached shock wave thus the majority of the heat generated was dissipated back into the atmosphere leaving only radiated heat to contend with, unlike a sharply pointed body where the shock wave was attached to the tip, causing heat transfer and destruction of the body. Additionally the heat reaching the blunt body would be more evenly distributed, preventing hot spots more prone to burn through.

Allen and Eggers demonstrated that the maximum deceleration encountered by a reentry vehicle was a function of the angle of reentry as well as velocity and independent of the shape, size and mass or drag coefficient. The importance of shape was the amount of heat that was absorbed by the reentry vehicle. A team of Ames researchers led by Eggers and including Fred Hansen and Bernard Cunningham published a method for predicting heat transfer to blunt bodies in 1958 though the work was done and in use much earlier but not published for six years due to classification issues. (7)

In order to reach targets 4,000 to 6,000 nautical miles away, ballistic missiles would need to be accelerated to speeds of up to approximately Mach 20 (15,223 miles per hour, just short of orbital velocity), 10 times the speed of a high-powered rifle bullet. (8) Reentry into the atmosphere at these speeds would generate a shock wave in which the atmosphere is heated to many thousands of degrees, even approaching 12,000 F, which exceeded the melting point of tungsten, the metallic element with the highest known melting point, 6,116 degrees Fahrenheit. (9) At this temperature the air plasma is also highly chemically reactive. There is a transport of heat by mass conduction from the air plasma to the vehicle surface which is dependent on both the temperature and density of the air in the plasma. At high altitudes where the air density is low, the mass transport of heat is low, in spite of the very high shock wave temperature. Conversely, at lower altitudes, the higher density plasma results in a higher heat flux for equal reentry vehicle velocities(Figure 2). (10)

Before discussing individual test and operational reentry vehicles, a brief discussion of testing methods, both for ground and flight is necessary.

Reentry Research Tools

Hypersonic Wind Tunnels

While the history of the military use of ballistic missiles rightly starts with the development of the A-4 (V-2) missile, perhaps just as important was the discovery by Allied troops of two highly advanced wind tunnel facilities at Peenemunde in the summer of 1945. One had apparently been in operation, a small diameter (1.2 foot) super-supersonic wind tunnel for intermittent use up to Mach 5 and a larger diameter (3.3 foot) continuous flow super-supersonic wind tunnel designed for speeds up to Mach 10.

In 1945 the first hypersonic wind tunnel in the United States was proposed by John Becker at Langley. Design difficulties and a perceived lack of urgency by NACA and Langley administrators delayed the construction for over a year but on November 26,1947, the first tests were successfully run at Mach 6.9. (11) Eggers at Ames, proposed a continuous flow hypersonic tunnel and it was completed in 1950. Between these two facilities, hypersonic research began in earnest, mainly focusing on aerodynamic issues directed towards supersonic aircraft research.

By 1955, the three major ballistic missile programs, the Air Force Thor (IRBM) and Atlas (ICBM) and the Army Jupiter (IRBM), made reentry vehicle research a high national priority. Two flight regimes required detailed study. The 1,500 nautical-mile IRBM Thor and Jupiter warhead reentry speed would be nearly 15,000 feet per second while the 5,000 nautical mile range ICBM would be nearly 25,000 feet per second. (12) Basic ballistic shapes, along the lines suggested by Allen and Eggers were tested up to the Mach 7-10 capabilities of the early hypersonic wind tunnels, confirming their theoretical results. However, the limitations in run times and temperatures, as well as atmospheric densities, soon illustrated the need for additional testing facilities.

Shock Tubes

The first shock tube was built in France in 1899 by Vielle to study flame fronts and propagation speeds resulting from explosions. (13) The concept languished until 1946 when Payman and Shepard in Britain published a thorough description of the design and use of shock tubes in studying explosions in mines. (14)

There are many variations of shock tube design but all share a basic two chamber concept. The first chamber is separated from the second with a burst diaphragm calculated to burst when the gas in the first chamber is compressed to a predetermined value. Since 1949, shock tubes have been used to augment aerodynamic studies using hypersonic wind tunnels, in particular the use by the mid-1950's was focused on reentry vehicle design and material selection since speeds greater than Mach 10 could easily be achieved, as well as much higher temperatures. The major drawback was the limited duration of test conditions. (15) Both Ames and Langley's Wallops Island Flight Test Range utilized shock tubes for reentry vehicle research. (16)

Avco Corporation learned of the shock tube work of Arthur Kantrowitz at Cornell University's School of Aeronautical Engineering funded by the Naval Ordnance Laboratory. Kantrowitz ran test models of the Mark 4 reentry vehicle that Avco was developing as a back-up for the General Electric Mark 3 for Atlas and for use as the primary reentry vehicle for the Titan I. In 1956 he left Cornell to head up the Avco Everett Research Laboratory where he led development of the ablative materials for the final Mark 4 design as well as for the Minuteman Mark 5 and Mark 11 reentry vehicles. (17)

Light-Gas Gun

The two stage light-gas gun was invented in 1948 by E.J. Workman at the New Mexico Institute of Mining as a method to dramatically increase projectile velocity. Despite the impressive German and Russian developments in artillery during World War II, perhaps the most famous of which was the German Tiger Tank 88 mm gun, projectile velocities remained at an upper limit of 9,000 feet/second.

The basic concept of the light-gas gun was to replace the gaseous byproducts of conventional gun powders which propelled the projectile, with a column of hydrogen or helium. A standard gunpowder cartridge was used to fire a plug down a barrel filled with helium or hydrogen (hence the term light-gas) which would compress to the bursting point a diaphragm immediately behind the actual test projectile. When the diaphragm burst, compressed light gas would propel the projectile down a second barreling allowing far greater velocities to be achieved since the molecular weight of the propellant gas would now be approximately l/8th of that of the water, carbon dioxide and nitrogen byproducts of gunpowder combustion (4 g/mole for helium versus approximately 30 g/mole) (Figure 3).

Workman's research group received funding from the Army Ballistic Research Laboratory (BRL) and proved the concept, reaching a velocity of 9,800 feet per second and quickly extending it to nearly 14,000 feet per second. The results caught the attention of the BRL managers, the device declared classified and removed, with all of the associated equipment, to the BRL facilities. Work did not continue at BRL for reasons that are not clear.

With the need for a relatively inexpensive method to "flight" test small models of proposed Atlas and Thor reentry vehicles, in the mid-1950's the light-gas gun concept was given new life via contractors and universities as well as researchers at both Langley and Ames. Velocities were soon extended beyond 25,000 feet per second. (18)

Atmospheric Entry Simulator

In early 1955, Eggers at Ames pondered the idea of simulating reentry through the varying densities of the upper and lower atmosphere. Could a method be found for launching a test article at reentry speeds into a test chamber that could simulate the gradual increase in atmospheric density which was the most problematic for the thermal stress of reentry? A light-gas gun could be used for launching the test article as their development had progressed to provide reentry velocities but how to simulate the atmosphere at 100,000 feet where most of the aerodynamic heating takes place? The necessary 100-fold variation in atmospheric density in this part of the reentry envelope might be achieved using components of a supersonic wind tunnel, the settling chamber and the exit portion of a Mach 5 supersonic nozzle. Eggers reasoned that the light-gas gun could be used to fire a small scale reentry vehicle model into the Mach 5 supersonic nozzle and then caught for detailed examination. The result was a small prototype Atmospheric Entry Simulator (AES) which was built in 1956, and successfully tested in 1957, evolving into a larger version in 1957. (19) This large AES was used successfully in exploratory work on blunt body copper heatsink designs meant for use on the shorter range and substantially lower heat regime IRBM missiles with reentry speeds of 15,000 feet per second. (20)

Arc Jet

Major drawbacks to the methods already addressed was still the relatively short duration of velocities, temperatures and inability to reach the higher temperatures of reentry in a continuous flow wind tunnel. After investigating several possibilities, the solution appeared to be the use of an arc-jet heater. Research at Ames began in 1956 and resulted, six years later, in the Gas Dynamics Laboratory devoted to further arc-jet development for use in stand-alone testing of ablation materials. While arc-jet wind tunnels are used to study reentry phenomena in a step-wise manner, they are unable to simulate conditions of a constantly descending reentry vehicle. (21) Several different types of arc-jet heaters, including subsonic air arc jet heaters and arc-jet radiant heaters are also used outside of a wind tunnel to study the ablative properties of materials. The arc-jet, with its more easily managed test conditions as well as longer test duration times, along with the fact that the test model was held in place, eventually replaced the AES for study of ablative materials at Ames.

Avco Corporation's Everett Research Laboratory and General Electric's Missile and Space Vehicle Division, amongst other labs, also employed variations of the arc-jet in their research and development of ablative materials for use on reentry vehicles. In 1958 James Fay, from the Massachusetts Institute of Technology and Avco's Frederick Riddell published a theory that allowed calculation of boundary layer conditions in high speed flight: (22)
The boundary-layer equations are developed in general for the case of
very high speed flight where the external flow I in a dissociated
state. In particular the effects of diffusion and of atom recombination
in the boundary layer are included. It is shown that at the stagnation
point the equations can be reduced exactly to a set of nonlinear
ordinary differential equations even when the chemical reactions
proceed so slowly that the boundary layer is not in thermochemical
equilibrium.


P.H. Rose and W. I. Stark at Avco published a paper at the same time comparing the theory against shock tube experimental results: (23)
Simulation of flight stagnation conditions at velocities up to
satellite velocity of 26,000 feet per second is shown to be possible in
shock tubes and data has been obtained over a large altitude range at
these velocities.


These two papers extended that of Lester Lees published in 1956 which had been found to underestimate by as much as 30 percent heat transfer rates at the reentry vehicle tip. (24) Now reentry vehicle researchers had both experimental and theoretical methods for evaluating ICBM reentry vehicle materials and possible designs.

Rocket Motor Exhaust

Development of the Jupiter IRBM reentry vehicle took place at the Army Ballistic Missile Agency (ABMA) facilities at the Redstone Arsenal, Huntsville, Alabama. Researchers there used the exhaust from a number of different liquid rocket engines to test candidate jet vane materials to replace the troublesome graphite vanes used in the V-2. (25)

Solutions to the "Reentry Problem"

Theodore von Karman, perhaps the leading aerodynamics expert of his time, described what he called "the reentry problem" at a symposium in Berkeley, California, June 1956. Reentering the atmosphere at speeds of Mach 12-20 was "perhaps one of the most difficult problems one can imagine. . . a challenge to the best brains working in these domains of modern astrophysics." (26) While the workers at Ames, Langley and other facilities had partially met the challenge via theoretical calculations about vehicle shape which led to the design of testing facilities, what was the solution to the remaining aspect, taming the thermal load encountered at these high speeds?

Four categories of cooling were considered: a) radiant cooling via emittance from the vehicle surface, b) solid heatsinks which would have sufficient mass to absorb the heat and protect the payload, c) transpiration and film cooling which would cause heat removal by material phase change, d) ablation which would allow heat dissipation via the many protective processes associated with surface removal.

Each of the four options had specific environments where they were most effective. Radiant cooling was best for long duration reentry environments where heat load was relatively low and constant and in practice worked best at temperatures below 2,000 F. Solid heatsinks could accommodate higher temperatures as long as the heating rate was not so rapid as to melt the material. Additional large structural mass was necessary to store the heat and protect the payload. Transpiration and film cooling would be able to work over a wide thermal environment but were mechanically complicated which might reveal hidden reliability issues. Ablation worked well for short duration, high temperature environments, the question was one of which materials to select and how to test them. (27) Only two of these concepts, heatsink and ablation, were used in research and operational reentry vehicles.

A key description of a reentry vehicle is its ballistic coefficient, beta (p). is defined as W/([C.sub.d] x A), where W is the weight of the reentry vehicle, [C.sub.d] is the coefficient of drag and A is the cross-sectional area. With reentry vehicle weight being held constant, reentry vehicles with a low [beta] (high coefficient of drag and cross-sectional area, and thus high air resistance) decelerate at a relatively high altitude, where the density of the atmosphere is low and heat fluxes are lower but reentry times are longer, facilitating radar detection while simultaneously resulting in decreased accuracy. Medium [beta] vehicles decelerate at a medium altitude with higher heat fluxes but shorter detection times and increased accuracy. High [beta] vehicles decelerate at much lower altitudes, encountering much denser air and hence higher heat fluxes but for a shorter time, allowing less time for radar detection and also greatest accuracy. Obviously these considerations were critical to mission requirements but were constrained by both the materials and testing facilities available at the time.

The First Generation - Heatsink

The work of Allen and Eggers had clearly shown the importance of selecting a relatively blunt nose shape for ballistic missile reentry vehicles to minimize aerodynamic heating. There was still an enormous amount of heat to be dealt with and this meant selecting the best temperature-resistant and high strength materials. Allen and Eggers research showed that most of the aerodynamic heating would be outside the boundary layer and not in direct contact with the reentry vehicle provided the boundary layer remained laminar. A considerable amount of radiative heat still had to be dissipated. Since radiation varies as the fourth power of the temperature, it was likely that the reentry vehicle would not be an efficient radiator with the result that surface temperature would rise beyond either the structural stability of then currently available materials or the tolerance level of the enclosed equipment, i.e., fusing and actual warhead. Heavily influenced by Allen and Eggers seminal work in conjunction with the paucity of high temperature stable materials, the first choice for reentry vehicle heat control was the heatsink concept. Both the Navy and Air Force elected to use the heatsink concept for their first generation reentry vehicles. The Air Force program is known in greater detail but both are discussed next because the Navy had a novel approach to reentry vehicle design (Table 1).

Navy

Mark 1

As with Thor and Atlas, the reentry vehicle (the Navy used the term reentry body but reentry vehicle is used here for consistency) needs for Jupiter-S (progenitor to Polaris) coincided with the viability of the heatsink concept since ablative material research was still relatively new in 1955 had not progressed far enough.(Figure 4).

The Navy quickly moved from the Jupiter-S program to Polaris. Due to weight constraints imposed by the Polaris missile solid engine performance, the reentry vehicle/warhead combination had to be much lighter than the Jupiter payload with a goal of a nearly seventy percent reduction to 1,000 pounds, at most. Consequently, the Navy was focusing, unlike the Air Force and Army reentry vehicle designs, on a reentry vehicle that did not encase the warhead. Instead, the warhead would ideally be an integral part of the design. (28)

On December 21, 1956, the Navy Bureau of Ordnance asked the NACA to study reentry body shapes for use in the new Polaris IRBM program. Just one day earlier a flight test at Wallops Island had shown that using a flat-faced cylinder sub-scale model made of copper, the design could survive reentry speeds of Mach 13.9. Additionally the superiority of copper over Inconel-X was also proven. (29) Earlier work in 1956 with five flat-face and one hemispherical shape at Mach 2 illustrated the potential for blunt nose shapes with the flat-face shapes showing substantially reduced heat transfer (Figure 5). (30) In mid-1958, a feasibility study was published by James R. Hall and Benjamin J. Garland of Wallops Island Pilotless Aircraft Research Division. Two possible flat-faced cylindrical shapes with flared ends were evaluated. Their calculations showed that a flat-faced cylindrical shape with a flared afterbody was possible and if made of beryllium (the cylindrical part was assume to be the outer casing of the warhead) the resultant vehicle would be 134 pounds lighter than if composed of copper (Figure 6). Soon backed up by additional ground and flight testing, the Polaris reentry vehicle shape was close at hand. (31)

The Navy used flight test systems at Cape Canaveral and Wallops Island. At Cape Canaveral a modified Air Force X-17 rocket was used in a four flight FTV-3 series to evaluate reentry vehicle shapes and materials. These flights took place from July 17, 1957 to October 1, 1957 and all were successful. Three flight test programs were conducted at Wallops Island in support of the Polaris reentry vehicle development, with fifteen flights between March 1958 and August 1959. (32)

The addition of a hydrodynamic faring which covered the flat nose shape and was ejected once the missile began flight was all that was left to complete the shape of the Mark 1. At the September 26, 1957 meeting of the Special Projects Office Steering Task Group, evaluation of heatsink materials had narrowed down the W47 nose cap material to beryllium or copper. Knemeyer at China Lake had read a RAND study on reentry heat shield materials and noticed that beryllium was an excellent candidate from a heat shield standpoint as well as the fact that the warhead casing was also made of beryllium. (33) The decision to use beryllium, at the time not a commonly used metal or readily available in the United States and which had only been used in alloy with copper, was somewhat controversial. The controversy stemmed from the issue that the Atomic Energy Commission (AEC) was being told by the Navy which material should be used for the casing of the warhead. The AEC resisted the suggestion at first but armed with the results of the Hall and Garland study, the Navy persisted and prevailed. (Figure 7). (34)

Air Force

On January 24, 1955, the Air Force and Lockheed Aircraft Corporation (Lockheed) signed a letter of intent authorizing Lockheed to develop and conduct a program into the design of reentry vehicles. At this time none of the currently available aerodynamic research facilities in the country could simulate the high thermal and velocity conditions of long range ballistic missiles. New techniques were becoming available but the conclusions reached from them needed to be confirmed with actual flight test data.

The result was the X-17, designed to achieve a reentry speed of Mach 15 and achieve a Reynolds number of 24 million (the Reynolds number is an indication of viscosity with a high value indicating viscosity is negligible) while measuring boundary layer conditions and the transition from laminar (desired) to turbulent (undesired) flow around the reentry vehicle. The Air Force reasoned that the X-17 would be able to provide the required data without waiting years for full-scale Atlas or Titan missiles to be ready while also being much less expensive. Sub-scale reentry vehicle shapes and material could be screened quickly and appropriate conversion of the data to full-scale models could be made. (35)

On February 17, 1955, representatives from the Western Development Division, Ramo-Wooldrigde and Lockheed visited the Langley facilities at Wallops Island where a few months earlier the first Mach 10 flight of the Langley Pilotless Aircraft Research Division (PARD) had taken place using a four stage solid propellant vehicle. Unlike the proposed X-17 flight profile which focused on high speed reentry, the PARD program Mach 10 speed had been reached at 86,000 feet with a coast up to 219 statute miles and a down range distance of 400 nautical miles. The X-17 program was described with the hope that the PARD program could be expanded to include the X-17 program. The Air Force schedule of a dozen flights at Mach 15 within a year was incompatible with the existing PARD programs but the Air Force decide to support the ongoing PARD programs by transferring some of the Sergeant rocket motors assigned to the X-17 program to Langley for use at Wallops Island. (36)

The X-17 was a three stage solid propellant missile designed to expose sub-scale re-entry shapes and materials to conditions of Mach 15 and a Reynolds number of 24 million. The program had four phases, using quarter- and half-scale rockets for development purposes and full-scale airframes for the research phase. For the full-scale rocket, 40.5 feet in length and weighing 12,000 pounds (8,500 pounds of propellant), the first stage was a single 31 inch diameter Sergeant motor, the second stage was a cluster of three Recruit motors 18.4 inches in diameter and, and the third stage a single Recruit motor, 9.72 inches in diameter. (37) The X-17 flight program began on May 23, 1955 using quarter-scale models, moving to half-scale on June 23, 1955 and the full-scale rocket on August 26, 1955, ending with the seventh full-scale flight on June 26, 1956. The fourth phase began on July 17, 1956 and ended on March 21, 1957 with only two failures out of thirty-six test flights. The two failures were caused by airframe problems and not propellant or staging issues, thus demonstrating the reliability, of multistage a solid propellant system. (38)

The flight profile emphasized the type of reentry conditions foreseen for ICBM reentry vehicles. The first stage propelled the airframe to 90,000 feet at burnout (Figure 8, 9). The missile then coasted to an altitude of 300,000 to 517,000 feet depending on the launch angle and vehicle weight. As the missile fell back to earth, the four fins on the first stage assured that the missile orientation was nose down. At an altitude of 90,000 to 70,000 feet, depending on the test objectives, a pressure probe initiated stage separation and ignition of the second stage along with activating a delayed signal for third stage ignition. At third stage burnout, speeds of Mach 11.2 to 14.5 were reached at 55,000 feet, again depending on launch angle. No effort was made to recover the reentry vehicle models, they lasted only long enough for telemetry on heating rates to be transmitted and often completely consumed. Of the 24 research phase flights, 18 were completely successful, one partial successful and five were failures. Blunt, hemisphere and cubic paraboloid reentry vehicle nose shapes were flown with six flights each for the General Electric and Avco blunt heatsink shapes being developed for the Atlas and Thor programs. (39)

Mark 2

The smaller the radius of the nose cone, the higher the temperature generated by atmospheric friction. By 1955, the scientists at the Army Ballistic Missile Agency (ABMA) had demonstrated to their satisfaction that the ablation method was the obvious direction to pursue, but the Air Force had opted for the more conservative approach of the heatsink method. If an ICBM was to be developed in a timely manner, to the Air Force way of thinking there was no other option but to go to a large radius, low [beta] reentry vehicle, the heatsink approach.

Much earlier work by Convair, the Atlas prime contractor, had pointed towards transpiration cooling for the reentry vehicle. The resultant weight, approximately 7,000 to 8,000 pounds, necessitated the original five engine design. Convair wanted to use a ceramic reentry vehicle, possibly due to the Army's work on Jupiter but at the time fabrication techniques for this large a vehicle were not available. When Ramo-Wooldridge (R-W) became the systems engineering contractor for the Western Development Division in 1954, they took a systems approach to reentry vehicle development. A blunt heatsink reentry vehicle design was well within the laboratory investigation abilities at that time. On December 22, 1954, R-W, Sandia Corporation and the Atomic Energy Commission agreed that the proposed Convair reentry vehicle weight could be cut in half and still provide space for the one megaton yield warhead the Air Force required. The decrease in reentry vehicle weight, combined with a new 2,000 pound warhead, meant that the overall weight of the missile could decrease from 460,000 pounds to 260,000 pounds and the propulsion unit reduced from five to three engines. (40)

General Electric (primary contractor) and Avco (backup contractor) were awarded an Air Force contract in 1955, to design, develop and manufacture a heatsink reentry vehicle for use on the Atlas ICBM. In this design, the heat of reentry was conducted from the surface to a mass of high heat capacity material rapidly enough to keep the surface temperature below the melting point of the shield material. Additionally, the mass of the heatsink absorbed the heat and prevented the payload from suffering thermal stress. The Air Force's scientific advisors concurred with the heatsink decision and the General Electric "froze" the design in terms of the warhead dimensions and heatsink method on 5 September 1956. (41) When the Air Force was assigned the Thor IRBM program, the Atlas reentry vehicle design was shifted to accommodate both missiles, saving development costs since an reentry vehicle designed for ICBM conditions would easily withstand the less strenuous conditions of IRBM reentry. (42)

Work by Jackson Stadler at Ames in 1957, evaluated copper, Inconel-X, graphite and beryllium for use in heatsink reentry vehicles. Copper represented an example of an easily machined material with high thermal conductivity but relatively low melting point, 1,984 F. Inconel-X was an example of refractory metal (resistant to heat and wear), but had low thermal conductivity and a 1,200 F melting point as well as being difficult to machine. Beryllium was an example of a lightweight metal with high strength, excellent thermal conductivity, a melting point of 2,348 F, but was difficult to machine as well as being hard to supply in quantity at the time. Additionally the dust generated by machining was highly toxic. Graphite was an example of a semi-metal with high thermal conductivity and highest melting point, 6,442 F, and high sublimation temperature. Stadler's evaluation included: a) thickness of material to prevent melting or sublimation at the surface, b) weight of material thick enough to meet (a), and c) determining thermal stress due to temperature gradients in the material.

Stadler concluded copper was a likely candidate due to the mass of material being resistant to thermal shock (weight was a drawback) and protection from oxidation would be needed. Inconel-X was "completely unsatisfactory" due to the low thermal conductivity causing melting to occur early in reentry and little heat was transferred to the interior. Graphite was superior to copper from the standpoint of weight, requiring 1/24 the weight of copper for equivalent protection. Unfortunately it would require to be coated which would inhibit exploitation of the high sublimation temperature. Beryllium was attractive due to a higher melting point then copper and being much lighter, l/6th the equivalent weight of copper, but it was brittle and difficult to form in large pieces at that time. (43)

For the General Electric Mark 2 design copper was selected due to its ease of machining, high heat capacity and high thermal conductivity which meant the heat generated would be rapidly absorbed into the mass of copper and not cause melting at the surface. Avco scientists pursued the use of beryllium and were successful in creating a Mark 2 reentry vehicle but it was too late as ablation took over as the method of choice. The techniques developed were used to fabricate early research and development beryllium heat shields for Project Mercury. (44)

Work by Katherine C. Speegle at Wallops Island's preflight jet test facility in 1957, investigated the best shape for the nose and the compartment that would contain the warhead. Six blunt nose shapes with identical afterbodies were tested at Mach 2.0 velocities. The results showed that the selected truncated cone afterbody was completely surrounded by the separated flow region, meaning heating would be acceptable. (45) The final design was known as a blunt conic sphere. The Mark 2 had a maximum diameter of 63.6 inches and was 60 inches in length, weighing nearly 2,000 pounds (Figure 10). (46) The blunt conic-sphere was inherently unstable and prone to oscillations causing turbulent flow to develop on the nose of the vehicle so a trajectory control system was incorporated to provide rate damping of the oscillations as well as impart spin to increase accuracy. A Mark 1 reentry vehicle was initially developed as a flight article but due to changes in missile flight schedules was not flown and instead used for development fit testing and as a flight reserve article. (47) The surface of the Mark 2 was coated with a thin layer of nickel to decrease radiative heating and was highly polished to prevent localized hot spots. (48)

The X-17 program had demonstrated that an ionized air layer surrounding the vehicle during the highest temperature period of reentry caused a telemetry black-out. For full-scale flight testing of the Mark 2, General Electric engineers developed a buoyant data capsule. The capsules were 18-inch spheres made from two hollow hemispheres of polyurethane foam which housed a tape recorder, radio beacon, battery pack, dye pack and SOFAR (sound fixing and ranging) device for locating the capsule. The bottom half of the capsule was coated with shark repellent after a test capsule was recovered with a shark bite mark. The capsule was attached to a small rocket to boost it free of the reentry vehicle. The urethane sphere was encapsulated in an ablative shell which shattered on impact (40,000 g's), releasing the buoyant capsule. Contact with salt water triggered the release of dye, the SOFAR device and the radio beacon. (49)

The Atlas Mark 2 flight test program began on July 19, 1958 and ended on December 19, 1959, a total of seventeen flights; seven Atlas B, four Atlas C and six Atlas D, nine were successful flights. The Thor Mark 2 flight test program began on November 5, 1958 and ended on December 17, 1959, a total of twenty-eight flights, with twenty-four successful. Details on Mark 2 reentry vehicle performance on these flights remains classified. The Mark 2 Mod 4 operational warhead weighed 3,500 pounds of which 1,600 pounds was warhead weight and was only deployed on Atlas D gantry sites at Vandenberg AFB from 1959 to 1964 and on Thor missiles in England from 1959 to 1963. (50) (Figure 11).

The Second Generation - Ablative

The first to actually describe ablation was Dr. Robert H. Goddard in 1920: (51)
In the case of meteors, which enter the atmosphere with speeds as high
as 30 miles per second, the interior of the meteors remains cold, and
the erosion is due, to a large extent, to chipping or cracking of the
suddenly heated surface. For this reason, if the outer surface of the
apparatus were to consist of layers of a very infusible hard substance
with layers of a poor heat conductor between, the surface would not be
eroded to any considerable extent, especially as the velocity of the
apparatus would not be nearly so great as that of the average meteor.


The process of ablation during reentry is described as follows: (52)
As heating progresses, the outer layer of polymer may become viscous
and then begins to degrade, producing a foaming carbonaceous mass and
ultimately a porous carbon char. The char is a thermal insulation; the
interior is cooled by volatile material percolating through it from the
decomposing polymer. During the percolation process, the volatile
materials are heated to very high temperatures with decomposition to
low molecular weight species, which are injected into the boundary
layer of air. This mass injection creates a blocking action, which
reduces the heat transfer in the material. Thus, a char-forming resin
acts as a self-regulating ablation radiator, providing thermal
protection through transpirational cooling and insulation. The
efficiency, in terms of heat absorbed per weight of material lost, is
about 40 times that of the earlier copper heatsink design.


Army

Jupiter

Ablation provided thermal protection for the Jupiter reentry vehicle. Earlier work had shown the transpirational cooling approach, while it worked, required complicated plumbing that would likely be hard to support in the field. The heatsink concept would work but was determined to be too heavy. The ablative approach came from a fortuitous result of research begun in 1953, investigating materials to replace graphite for jet vane application during the development of the Redstone missile (jet vanes were used for directional control instead of gimbaling the engine). The trouble was one of quality control because while a source of the right grade of material was found, the manufacturer's poor quality control meant that only twenty-five percent of the jet vanes were acceptable. In an attempt to find a replacement, researchers tested several materials including a jet vane made of commercial grade fiberglass-reinforced melamine. Exposure to the Redstone rocket motor exhaust eroded the vane as expected but much to the surprise of the researchers, one-quarter inch beneath the surface the material was not only undisturbed but the embedded thermocouples revealed no heating had taken place. While the tested material was not used as a jet vane, the ABMA researchers skipped past the heatsink concept and went straight to ablative reentry vehicle materials. (53) Ceramic material was also carefully evaluated and found to be too sensitive to thermal shock at that time though sufficient work was done with a method of ceramic manufacture called slip forming to successfully fabricate the necessary shape.

Scientists at ABMA estimated the weight of five candidate materials: Refrasil-phenolic, fiberglgass-melamine, unfired ceramic, beryllium and copper to provide thermal protection for a proposed heat shield design. Refrasil, fiberglass-melamine and ceramic were found to be the materials of choice. An expedient method for evaluating candidate materials was to expose flat plates of the material to rocket exhaust at a heat flux of 100 BTU/[ft.sup.2]-sec and a velocity of 6,700 feet per second. The plates were four inches square and tilted at a 45 degree angle in the exhaust stream. Further research in resin based ablative materials revealed that asbestos reinforced phenolic resin would be the best overall material for the Jupiter reentry vehicle environment. After initial evaluation of the plate material, reentry vehicle shapes were tested both with the rocket exhaust technique and via shock tube studies by Arthur Kantrowitz at Cornell University. (54) Using a variety of rocket motors, researchers were able to simulate heating rates up to 2,500 BTU/[ft.sup.2]-sec. Transonic wind tunnel tests of a half-scale model Jupiter reentry vehicle were conducted at the Air Force's Arnold Engineering Development Center, Arnold Air Force Base in June 1957 and at the hypersonic test facilities of the Naval Ordnance Laboratory, White Oak, Maryland in September 1957, confirming the full-scale nose cone design. (55)

For flight testing of the one-third scale Jupiter reentry vehicle designs, the Army's Redstone tactical ballistic missile was modified into a three stage booster. The first stage had an elongated fuel tank and used a more powerful fuel called Hydne (unsymmetrical dimethyl hydrazine). The forward section of the first stage was strengthened to support the new upper stages. The second stage was made up of a cluster of eleven scaled-down Sergeant solid propellant missiles, six inches in diameter, housed in a cylindrical fairing called the "tub." The third stage was located in the center of the second stage and made up of three additional scaled down Sergeant missiles. Atop the third stage was a 300-pound, l/3rd-scale ablative (1/10th surface area) reentry vehicle composed of a welded steel shell supporting the heat shield. While fabrication techniques were being perfected for the resin-asbestos material, a heat shield made of layered disks of melamine, a commercially available laminated fiberglass-resin was flown first. The tub was spun up by electric motors at launch to provide ballistic stability. The resulting vehicle was called Jupiter-C (Jupiter Composite) and now had a range of over 1,500 nautical miles with an apogee of over 175 nautical miles. (56) (Figure 12)

Only three of a scheduled of thirteen flights were necessary for the Jupiter-C program. The first launch was on September 20, 1956, Jupiter C Missile RS-27, with the missile reaching 600 nautical miles in altitude and a speed of Mach 18. This was a test of the modified propulsion and staging system and was successful. The second flight, Jupiter C Missile RS-34, was launched on May 15, 1957. The missile pitched up at 134 seconds into flight so while the planned range was not reached and the reentry vehicle was not recovered, telemetry indicated that the fiberglass melamine ablative material had functioned as expected. The first sub-scale operational Jupiter reentry using a phenolic resin asbestos ablative material was flown on Jupiter C Missile RS-40, August 8, 1957. The booster and high-speed upper stages worked well. Failure of the reentry vehicle to separate from the third stage changed the reentry trajectory, reducing the angle of attack at the point of maximum heating. Nonetheless, the reentry vehicle traveled 1,168 nautical miles, achieving a velocity of 13,000 feet per second and withstanding a temperature of over 2,000 degrees F, conditions similar to those expected for an IRBM reentry vehicle. While the reentry vehicle did not separate as planned, the heat of reentry melted the magnesium ring of the separation system and the recovery system deployed successfully. Analysis of the ablative covering showed only a one and a half percent loss (the reentry vehicle was displayed in President Eisenhower's office and is in storage at the National Air and Space Museum in Washington, D.C.) Ablation technology had been proven with the ultimate test, full IRBM range and velocity. (57)

Full-scale Jupiter reentry vehicles were successfully recovered on three flights; Jupiter Missile AM-5, launched on May 18, 1958, the first recovery of an IRBM reentry vehicle; Jupiter Missile AM-6, July 17, 1958, which also carried a lightweight high explosive warhead; and Jupiter Missile AM-18, May 28, 1959, which carried two monkeys, Able and Baker, which survived unharmed. While the reentry vehicle flown on AM-5 showed an ablation depth of three-eighths inch at the greatest point of loss, the remaining flights showed considerably less, validating the ablative concepts of the sub-scale model flown and recovered earlier (Figure 12). (58)

The deployed reentry vehicle, built by Goodyear Aircraft Corporation, was an hermetically sealed conical aluminum shell with a twelve and a half-inch radius spherical tip attached to a cone frustrum with a base 65 inches in diameter and an overall length of nine feet. The molded nose cap was composed of thirty percent, by weight, phenolic resin with seventy percent Type E glass; the frustrum material was a layer of a mixture of forty-five percent phenolic resin and fifty-five percent Chrysotile asbestos. (59) A key design feature, also found in other reentry vehicle designs, was a convex dish shaped aft cover which conferred the ability to recover from any attitude to the correct reentry alignment. The ablative material was much thinner than the sub-scale fiberglass melamine heatshield. (Figure 13). The complete reentry vehicle with warhead, weighed 2,617 pounds, the W49 weapon weighed 1,600 pounds. (60)

Air Force

Atlas and Titan I

The Air Force was not new to the concept of ablation. Indeed the two contractors selected in 1955, to develop the Atlas reentry vehicle, General Electric and Avco Corporation, were directed to look at all methods for solving the reentry problem. Wright-Patterson's Air Research and Development Command were also evaluating ablation materials as was Langley Aeronautical Laboratory and Ames Research Center. The decision to work with the heatsink concept had stemmed from recommendations of a number of scientific advisory committees and panels. On June 16, 1953, the Department of Defense Study Group on Guided Missiles, better known as the Teapot Committee, had been created to evaluate the status of guided missile development by the Air Force. On February 16, 1954, the committee submitted its report. It recommended that the reentry problem be reinvestigated as Convair's approach (transpirational cooling) was insufficiently broad. (61)

On August 31, 1957, in the 21st Monthly Report on Progress of ICBM and IRBM Programs, a shift in reentry vehicle design was noted. While the heatsink design for Atlas and Thor was sufficient, developments in materials and testing capability indicated that ablation reentry vehicles could have ballistic coefficients five to eight times greater than the Mark 2 heatsink which would lead to greater accuracy and decreased vulnerability. Dispersion caused by wind would also be greatly decreased, the error due to wind was calculated to be approximately 500 feet in CEP (circular error probable, a circle within which fifty percent of the reentry vehicles impacted) at a 5,500 nautical mile range. (62)

On August 28, 1958, after only two Atlas B flights with the Mark 2 and just before the start of the Thor Mark 2 flight testing, almost exactly one year after the highly successful conclusion of the Army's Jupiter-C reentry test vehicle program, Brigadier General O.J. Ritland, Vice Commander of the Ballistic Missile Division, notified the Air Research and Development Command of the decision to reorient the ICBM reentry vehicle program from heatsink to ablative technology. The decision was based "recent developments aimed toward improving the solution to the ICBM reentry problem." The Mark 2 heatsink reentry vehicles would be supplied for all WS-315A (Thor) and early operational WS-107A-1 Atlas missiles at the two operational sites at Cooke Air Force Base (Cooke had not been renamed Vandenberg yet). All Avco work on heatsink development was to be discontinued. General Electric was now assigned development responsibility for a light weight second generation reentry vehicle capable of carrying a 1,600 pound warhead, and to be flight tested on the Series D Atlas missiles with deployment starting at Warren Air Force Base. This was the Mark 3. Avco was assigned responsibility for a heavy weight second generation reentry vehicle capable of carrying a 3,000 pound warhead to be flight tested on lot J Titan I missiles. This was the Mark 4. (63)

As early as 1956, plastics had been examined for use in the high temperature environment of ramjet engines. Researchers at the Marquardt Aircraft Company exposed model ramjet inlet cones made from three fiberglass reinforced plastics, Conolon 505 (phenolic), DC 2106 (silicone) and Vibrin 135 (polyester) for twenty minutes at temperatures up to 500 to 600 F at a speed of Mach 2. They found that all three materials showed little or no detrimental effects, concluding that reinforce plastics might have a role in missile development. (64)

Researchers at General Electric's Missile and Ordnance Systems Division in Philadelphia expanded on the Marquardt work by estimating a candidate ablative material's ability to absorb heat up to 8000 F under equilibrium conditions. The results showed that plastic materials had the highest theoretical heat absorbing capacities, more than twice that of beryllium. The more gas a material generated upon heating, again under equilibrium conditions, the better the material. Heat capacity and gas generation values were useful indicators but could not be used as guides in selection of materials because of the non-equilibrium conditions of the operational environment. When the material melted, the liquid would be swept away in the air stream, upsetting the thermal equilibrium. The higher the melting point and the more viscous the resulting liquid, the more optimal the thermal effect. Phenolic resin plastics were found to decompose slowly at high temperature and did not liquefy, instead forming gaseous byproducts and a char layer that protected to the base material. Exposure of phenolic-glass cloth with sixty-five percent resin to 12,000 F in a high temperature arc showed only 1.4 percent erosion; phenolic-Refrasil (Refrasil is the trade name for a high silica content glass) with forty-one percent resin only 2.1 percent erosion and phenolic-nylon cloth with fifty-seven percent resin only 1.0 percent. The organic reinforcement's lower erosion rate was due the organic fiber's lower thermal conductivity. Key variables were type of resin, orientation of the fibers, type of fiber and ratio of resin to fiber. Phenolic resins gave a higher yield of carbon char. Large variations in performance were found amongst the various suppliers. Orientation of the fibers had a significant effect on performance with random orientation giving the best results. At temperatures above 5,000 F amorphous silica fibers were superior to ordinary glass and organic fibers were found superior to glass fibers. Resin to fiber ratio optimization had somewhat counter intuitive results. Higher glass fiber content gave better mechanical properties but was slightly detrimental to thermal erosion above 5,000 F. At plasma jet temperatures, 12,000 F, higher resin content gave greatly improved performance. Clearly ablative materials had come of age for use in ICBM reentry vehicle heat shields. (65) The result was the General Electric's Mark 3 reentry vehicle deployed on Atlas D.

Avco Corporation began defense contract work in 1955, with the creation of the Avco Everett Research Laboratory. Victor Emanuel, president of Avco Corporation, knew of the work of Dr. Arthur Kantrowitz, a physicist working at Cornell University with shock tube experiments in the study of the hypersonic flight. Emanuel approached Kantrowitz with a proposal to come work at Avco and apply his theories towards the solution of the "reentry problem." Enticed by the prospect of a new, modern facility to be built for him, Kantrowitz agreed and the Avco Everett Research Laboratory was built. At the same time and undoubtedly due to Kantrowitz's presence, the lab's Research and Advanced Development Division won the backup contract for the Mark 2 heatsink reentry vehicle and was the primary contractor for a similar design for Titan I. Like General Electric, Avco was also studying and developing ablative as well as heatsink material. Unlike the engineers at General Electric who had studied ceramics and dismissed them as too difficult to work with compared to reinforced plastic resins, Avco engineers decided to pursue the use of ceramics for the nose section of the reentry vehicle where the heating was the most severe.

Expanding on the ceramics research by Georgia Institute of Technology and Battelle Memorial Institute for the Jupiter program, Avco researchers focused on solving the brittle fracture problem which was preventing the fabrication of the large and complicated reentry vehicle shapes light enough to be practical. The weight issue was the result of the amount of material needed to be structurally sound and not one of thermal protection efficacy. The decision was made not to search for new materials but rather to focus on new fabrication techniques. One solution investigated was the use of small ceramic tiles. This was rejected due to the thinness of the tiles and difficulty in assembling them on the curved nose section. The eventual solution was to use a metal honeycomb structure to hold small "pencils" of ceramic which did not easily fracture. By orienting the pieces in honeycomb cells at ninety degrees to the surface, optimum thermal protection and structural strength was obtained. In 1959, Avco's Research and Advanced Development Division announced the development of Avcoite, a magnesium honeycomb reinforced ceramic for use on the nose of the Mark 4 reentry vehicle originally destined for Titan I but which was also deployed on Atlas E and F (Figure 14). (66)

Flight Testing

Once the feasibility of ablative material had been experimentally determined, flight testing of sub-scale reentry vehicles began. The primary research and development flight testing for evaluating the early Air Force sub-scale and full-scale ablative ICBM reentry vehicles were the Thor-Able 0 and II, Atlas D and Titan I Lot J programs.

Thor-Able

The first in a series of ballistic missiles used for Air Force reentry vehicle development was the Thor-Able launch vehicle. Use of research and development flights of the Atlas ICBM was considered and rejected at this point as integration of reentry vehicle testing would interfere with the early development objectives. In October 1957, the Ballistic Missile Division and Space Technologies Laboratory began the design of the Advanced Reentry Test Vehicle (ARTV) that could be ready for use within six to eight months using existing hardware. The critical capability of the ARTV would be to reach ICBM reentry speeds of approximately 24,000 feet per second carrying a one-half scale reentry vehicle. A variety of possible test vehicle combinations were examined but only one that met the requirements of availability and performance; a Thor first stage and Vanguard second stage modified with eight spin rockets was configured by STL with autopilot and cutoff controls assembled from available Thor and Atlas components. (67)

Able RTV

The Thor-Able 0 program flight tested three General Electric reentry vehicle development models, designated Able RTVs. The RTV's were biconic-spheres 34 inches long and a base 38 inches in diameter, weighing 620 pounds and (Figure 15) fabricated with ablative material and flown from Cape Canaveral from April 23, 1958 to July 23, 1958. There was one failure due to booster malfunction and two partial successes. All three flights carried biomedical experiments with mice and while the two successes clearly demonstrated the efficacy of ablation at ICBM ranges and speeds, the reentry vehicles were not recovered as planned. The data provided by these tests helped determine how much the heat shield weight could be decreased and still be effective as well as verifying the superior performance of ablative materials compared to the heatsink materials. A description of the RTV series vehicle's ablative materials has proven elusive. (68)

Able RVX-1

For the Thor-Able II program, a modified Thor booster was used with its guidance package removed and the radio-inertial guidance system for the Titan I ICBM installed in the RVX-1 reentry vehicle. These six flights were designated as Precisely Guided Reentry Test Vehicle launches with two goals; evaluating the new guidance system which would also indicate the exact point of impact as well as continue to evaluate new ablative materials. (69)

Instead of using the recovery system from the Jupiter reentry vehicle test program which had been proven unsuccessful with the Thor-Able 0 flights, General Electric developed a more robust system to handle the much heavier RVX-1 vehicles. Additionally, the data capsule concept used in the Mark 2 program was used to record the telemetry during the flight and reentry phase when ionization phenomena prevented telemetry transmission.

General Electric provided the RVX-1 internal frame used to test both the General Electric and Avco Corporation ablative materials. The RVX-1 was a conic sphere flared-cylinder configuration (Figure 16), sixty-seven inches long with a cylinder diameter of fifteen inches and a flare diameter of twenty inches, and weighed 645 pounds. The flights began on January 23, 1959, starting with the RVX-1 carrying General Electric materials and alternating flights with Avco materials, and ended on June 11, 1959 with one failure, three partial successes and two complete successes. The General Electric RVX-1 tested three types of phenolic nylon ablative materials (phenolic nylon, phenolic glass and phenolic Refrasil) in sixty degree segments repeating every 180 degrees on the cylinder and flare. The nose was made of a thick layer of molded phenolic resin with one- inch squares of nylon cloth. (70)

The Avco RVX-1 vehicles (sixty-eight inches in length with a nose cap of eleven inches, a cylinder diameter of seventeen inches, cylinder length of thirty-nine inches, a flare length of eighteen inches and a flare base diameter of twenty-eight inches) had Avcoite on the nose and phenolic Refrasil tape covering the mid-section and flare. (71) On the April 8, 1959, the Avco RVX-1-5 was successfully flown 5,000 nautical miles down range with a maximum altitude of 764 miles and a reentry speed of 15,000 miles per hour (Figure 17). The nose cap easily withstood the heat of reentry as had the Refrasil material coating the cylinder and flare sections. The Avcoite ceramic had melted and flowed back asymmetrically a short distance down the cylindrical body as expected. Telemetry results indicated no effect on aerodynamic stability. Soon after recovery the nose cap was removed for further inspection and replaced with a mock-up due to security concerns. The RTV-1-5 is now in storage along with the removed nose cone at the National Air and Space Museum's Garber facility.

On May 21, 1959, the second General Electric RVX-1 flown was also successfully recovered, looking much the same as the Avco vehicle except that the reinforced phenolic-chopped nylon nose cap simply ablated and did not flow back along the cylinder. (72)

The RVX-1 flight program, even with the failures due to not recovering all of the vehicles (complete telemetry was obtained via the data capsules), further confirmed the maturity of ablative materials for use in high speed reentry as the RVX-1 vehicles were exposed to temperatures exceeding 12,000 F. The RVX-1 test vehicles were the direct progenitors of the General Electric Mark 3 (Atlas D) and Avco Mark 4 (Atlas E, F and Titan I) reentry vehicles. (73)

RVX-2 Series

By mid-1960 Atlas D missiles were available for use in the final phase of ablative material testing, flights of full-scale reentry vehicles at operational ranges and reentry speeds. The RVX-2 series involved tests of newer plastic ablative materials. Ranges flown varied from 4,400 nautical miles to the Ascension Island impact area to 6,400 nautical miles off the coast of Capetown, South Africa and further yet, 7,900 nautical miles to the Southern Atlantic off of the Prince Edward Islands. The reentry evaluation portion of the program commenced on March 8, 1960, and ended on January 23, 1961. (74)

RVX-2

Three General Electric RVX-2 reentry vehicles were flown to test a new type of ablative material, unrein-forced phenolic resin, General Electric Series 100, for the proposed Titan II Mark 6 reentry vehicle. (75) The RVX-2 was a conic-sphere configuration, twelve feet tall and five feet in diameter, weighing over 2,000 pounds, the largest reentry vehicle yet flown with what appears to a phenolic resin-chopped nylon nose cap and unreinforced phenolic resin side frustrum panels. The first two flights suffered guidance and booster failures; March 17, 1959 and March 18, 1959 respectively, but the last flight, on July 21, 1959, was successful and the reentry vehicle was recovered intact after a flight of 5,000 nautical miles.(Figure 18). Photographs of the recovered vehicle show a close resemblance to the General Electric Titan II Mark 6 reentry vehicle which also used these materials. (76)

RVX-2A

The RVX-2A program had three flights during the Atlas D test flight program, August 12, 1960, September 16, 1960 and October 13, 1960. The RVX-2A vehicle had the same dimensions as the RVX-2 and weighed slightly more than 2,700 pounds. The main difference between the two was the instrumentation, the RVX-2A was used for extensive scientific experiments beyond reentry. The eighteen experiments included black and white and color photography, live mice, radiation phenomena, reentry physics including transpirational cooling, electromagnetic propagation and fuel cell prototypes. A recovery system similar to that of the RVX-1 program was used on all of the flights with successful recovery on only the final flight. (77)

The General Electric portion of the RVX-2A program were the first and third flights, testing the General Science Century Series of unreinforced phenolic resin for use on the conic frustrum part of the conic sphere design for the Titan II Mark 6 reentry vehicle. Four formulations, GE Type 123,124 and 135 as well as GE Type 525 were used. General Electric researchers had discovered a radical departure from previous ablation research. Under laboratory test conditions simulating reentry, unreinforced phenolic resin formed several porous char layers one to two millimeters thick were formed in sequence. The first one quickly plugged up, was sloughed off by aerodynamic forces and was replaced instantly by the formation of a new char layer. Large amounts of pyrolysis gases that formed as the material degraded served to inhibit heat transfer from the very hot boundary layer to the ablating surface, greatly reducing the actual heating at the vehicle surface. These results greatly simplified the design of the large Mark 6 reentry vehicle and saved considerable weight. (78) The maximum internal temperatures reached in the two flights were 90 and 100 F, well below the 350 F expected. Nose cap ablation was greater than expected. Degradation of the Series 100 phenolic resin was comparable to that of nylon reinforced phenolic resin and was in agreement with computer modeling(Figure 19). (79)

Avco flew one RVX-2A flight on September 16. The nose cap was RaD 58D followed by a twenty-six-inch frustrum section of RaD 58B and 100 inches of tape wound Refrasil. Test plugs of Avocoat x3007 and RaD 58E were inserted at alternating ninety-degree intervals in the forward portion of the tape wound Refrasil section. The RaD 58E was a candidate material for the Minuteman missile reentry vehicle and the Avocoat was a proposed low temperature ablation for the boost phase of the Minuteman trajectory. Telemetry problems prevented transmission of thermal and ablation data. (80)

Mark 3

The Mark 3 reentry vehicle was designed for the Atlas F missile, as mentioned earlier, but deployed only on Atlas D. The Mark 3 was a direct descendant of the General Electric RVX-1 program. Measuring 114.8 inches in overall length, there were two Mark 3 shapes (see Figure 20). Both had the sphere-conical nose shape, 29.22 inches in length and a cylindrical mid-section 20.7 inches in diameter and 40.6 inches in length. The Mark 3 Mods I, IX and LA had a single biconic frustrum flare, 35.9 inches in diameter, that blended smoothly with the reentry vehicle adapter spacer atop the missile. The Mark 3 (Mods IB and IIB) had a biconic-2 shape with an second, wider flare at the base, 42.8 inches in diameter, resulting in a characteristic "skirt" conical ring slightly outwards above the spacer which was not modified to affect a more streamline appearance. The second flare aided in reentry stability by moving the aerodynamic center of pressure toward the rear of the vehicle.(Figure 20). Available photographic evidence indicates that the biconic-2 modification was the deployed version. The nose section was thermally protected by molded phenolic nylon, the mid-section and flare by tape wrapped phenolic nylon. (81)

Eleven full scale Mark 3 reentry vehicles were flight tested as part of the Atlas D research and development program from March 8, 1960 to January 23, 1961, with ten successful and one failure due to booster failure prior to launch. (82) The Mark 3 was deployed on Atlas D missiles from 1960-1965. (83) Mark 3 Mod 3 operational RV weighed 2,200 pounds of which 1,600 was the warhead. (84)

Mark 4

Unlike the General Electric Mark 3, the Avco Mark 4 design required additional experimental flights designated RVX-3, a 0.72 scale model and the 0.94 scale model RVX-4 due to modified Air Force requirements. The RVX-3 was flight tested on 5 Titan I C missile flights from December 12, 1959 to April 28, 1960. The RVX-4 was to have been the full-scale model but the diameter of the warhead was changed slightly, leading to the actual full-scale Mark 4. The RVX-4 was flight tested on one Atlas D and seven Titan I Lot G missiles. (85)

The Mark 4 was a sphere-cone-cylinder-biconic flare shape, 126.7 inches long, 33 inches in diameter at the cylindrical mid-section and 48 inches in diameter at the base of the flare. The Mark 4 flare varied from 7 to 22 degrees with two very small spin fins at the base of the flare. The nose cap was made of Avocite varying from 1.32 to 0.82 inches thick, the cylindrical body and flare protected by oblique tape wound Refrasil at 0.61 to 0.32 and 0.44 to 0.66 inches respectively; and the afterbody was protected with fiberglass. The Mark 4 with warhead weighed 3,800 pounds. (86) A second reference gives the operational Mark 4 as weighing 4,100 pounds of which 3,100 pounds was the warhead. (87)

The Mark 4 was flight tested on one Atlas D, seven Atlas E, seven Atlas F and twenty-eight Titan I Lot J and M missiles from October 11, 1960 to May 1, 1963. The Mark 4 was deployed on Atlas E and F and Titan I from 1962 to 196. (88) One Mark 4 was flown on Titan II during the Titan II research and development program. (89)

Mark 5

On January 13, 1958, in discussions within the Nose Cone Division of Space Systems, Ballistic Missile Division, a decision was made that initial design responsibility for the advance reentry vehicle for Minuteman would be Avco Corporation due to the heavy technical load already assigned to General Electric. On February 5, 1958, a letter was issued to Avco confirming the request for an advanced reentry vehicle design study which included design specifications. This was not a sole source contract for the reentry vehicle production, as with other reentry vehicles the contract would be a competitive one. (90)

With contractor bid proposals to be evaluated in late June, on May 28, 1958, the Nose Cone Division clarified the desire, previously discussed in the proposed preliminary operational concept of Minuteman dated April 8, 1958, for two reentry vehicle designs and two warheads. (91) One vehicle would have a weight of 790 pounds, including a 600 pound warhead for a range of 5,500 nautical miles; the second vehicle would have a weight of 550 pounds including a 350 pound warhead for a range of 6,500 nautical miles. The two designs would permit the preliminary operational plan target coverage from bases located in the southwest portion of the United States. The designs would be optimized for maximum range target coverage permitting each missile to fly to the maximum range estimated for the payload, negating the need for lesser range targeting for a given missile. Phase I and Phase II warhead feasibility studies had already been completed, permitting reentry vehicle dimensions to be established. Requests for proposals were issued to ten contractors for the reentry vehicle studies covering either or both of the reentry vehicles.

The Nose Cone Division explained the need for two reentry vehicles: (92)
There are several reasons which in our opinion make it imperative that
we continue with development of both vehicles. These relate primarily
to the warhead development itself. At the present time this country is
considering a moratorium on testing, the duration of this being
somewhat indeterminate. For this reason, AEC laboratories are
endeavoring to carry out during Hardtack all tests which appear to them
of importance in development of weapons for which requirements have
been stated. At the same time the AEC is attempting to get acceptance
by DOD of the concept of multi-use warheads. Under this concept which
has been favorable reception a weapon system requiring a warhead of a
particular weight will be forced to use an already available or planned
weapon which in some instances will have been developed for quite
different requirements. In our case for example the smaller warhead
will be that now under development for Nike-Zeus. Provided that the
Minuteman requirements are incorporated in the weapon design initially
which can be done if we establish our need for this weapon, there will
be no difficulty in obtaining maximum performance of the system (the
same is not true for the 600 pound Polaris warhead which must be
modified to a considerable degree to meet Minuteman requirements.) If
on the other hand we do not today establish a firm requirement for a
second, lighter warhead, it will be designed on the basis of Nike-Zeus
requirements and will be completely incompatible with the Minuteman
system in the event that we choose to use the second reentry vehicle at
some later date.

We feel emphatically that development must continue on both warheads
and hence both reentry vehicles since a requirement for one can not be
established without the other.


On July 20, 1958, AFBMD announced that Avco Corporation had been selected from a group of seven proposals (Aerophysics Allison, Avco Corporation, Ford Aeroneutronics, General Electrical, McDonnell, Republic Aviation and Douglas/Goodyear) to develop the two Minuteman reentry vehicles. Avco's role as the Mark 2 alternate source, as well as its research and development expertise with the new ablative materials gained from their work on alternatives to the Mark 2 were a key in their selection. The contract required development of a light and heavy reentry vehicle to accommodate two possible warheads designs weighing 350 and 600 pounds respectively, with warhead dimensions to be forthcoming. (93) The contract was formally awarded to Avco on September 19, 1958. (94)

The light version was cancelled December 4, 1958 to reduce costs (Avco was directed to continue studying the light version on a lower priority basis). The decision was based on the lower yield available for the light vehicle warhead as well as complications introduced into the missile test program by multiple combinations of reentry vehicles and the missile airframe. The result was a 790 pound reentry vehicle of which 600 pounds was due to the warhead. The larger reentry vehicle could also more easily accommodate changes in warhead dimensions. (95)

After nearly a year of indecision on the Minuteman warhead design on September 1, 1959, the Minuteman warhead was finally authorized. Avco's reentry vehicle sphere-cone-cylinder-flare design was based on the Mark 4 shape but was considerably smaller due to weight constraints. (Figure 21) Development work commenced on what was now called the Mark 5 reentry vehicle. Extensive wind tunnel and light gas gun evaluation of ablative material composition and thickness as well as studies of attitude control and structural design to withstand deceleration forces of twenty to fifty G's were undertaken. Flight test vehicles (Mark 5 Mod I) were in production by the end of 1960. Like the Mark 4, the Mark 5 had a nose cap of Avcoite, in this case Avcoite-1, bonded to the top of the cylindrical and flare sections which were machined out of a block of RaD-58B phenolic resin-Refrasil material. Reformulation of the ceramic material reduce the melting and flowing which occurred with the Mark 4 (Figure 22). The aft closure was configured to stabilize the RV during early reentry and was coated with Avcoat. The Mark 5 did not have an active attitude control system. It and the Mark 4 tumbled and upon entering the atmosphere small fins induced a stabilizing spin before the fins ablated early in reentry. (96)

The full-scale research and development flight test program began on February 1,1961, with the successful launch and flight of FTM-401, a fully configured Minuteman IA, from the Launch Complex 31A pad, Cape Canaveral Air Force Station, Florida. Two more pad launches took place, March 19, 1961 (failed) and July 27, 1961 (successful). Silo research and development launches at Cape Canaveral began on August 30, 1960 with a spectacular failure and ended on February 20, 1963 with six failures out of twenty-one launches. Mark 5 flight tests also utilized Atlas D (1), E (4) and F(3) missiles with one failure. The Atlas flight tests commenced on May 13, 1961 with a Mark 5 Mod I flown on an Atlas E and ended on July 31, 1963 with a successful Atlas D flight (Figure 23). (97) The Mark 5 Mod 5B weighed 300 pounds including SOFAR bomb. The Mark 5 was deployed on 150 Minuteman IA missiles beginning in 1962 and ending in 1969. (98)

Mark 11, 11A, 11B and 11C

In October 1960, the Department of Defense and the Atomic Energy Commission authorized development of an advanced version of the XW-56X1 warhead. In December 1960, the Air Force requested development of a lighter and higher yield warhead, designated the XW-59. One month later it was decided to have Avco develop a new reentry vehicle, the Mark 11, able to carry either of the new warhead designs and to be deployed starting with the second Minuteman wing, equipped with Minuteman 1B at Ellsworth Air Force Base. The Mark 11 series reentry vehicle had an operational requirement for a reduced radar cross section during the exoatmospheric portion of its trajectory. (99) ( Figure 24)

The Mark llseries, 11, 11A, B and C, had a somewhat similar size and shape to the Mark 5 but was slightly longer. Avcoite was not used in the nose section. RaD 58B was high silica content phenolic resin which was pressed into a block, machined to shape and then bonded to the reentry vehicle forecone. For the Mark 11, the body of the vehicle was made of RaD 60, a molded silica phenolic using chopped silica fibers which was machined to fit over an airframe made of fifty magnesium ribs that were covered with a spin formed magnesium skin for both the cylindrical and flare section (formed separately). The two assemblies were bonded with epoxy, the final machining completed, a radar cross section reducing mesh applied and final layer of Avcoat 2 applied. The pointed tip, a distinguishing feature of the Mark 11 series was made of glass fiber resin impregnated cloth molded on a mandrel and epoxied to the nose. It is was used to provide protection to the nose section radar cross section material from boost-phase heating. Once the Mark 11 entered the atmosphere, the nose and base fairing as well as the radar cross section reducing mesh were removed by ablation. At this point in reentry the vehicle was producing a highly ionized and readily detectable wake which was unavoidable. Unlike the Mark 4 and 5, the Mark 11 had small spin rockets to confer spin stabilization prior to reentry.

The Mark 11A, B and C had a different fabrication process from the Mark 11. The new aluminum frame was heavier than the Mark 11 magnesium frame but was stronger, a feature required for the nuclear hardening of the vehicle, a new operational requirement due to advances in the Soviet AntiBallistic Missile (ABM) system in the process of development. The flare, cylindrical body and nose cap frames were bolted together and then the heatshield applied using Oblique Tape Wound Refrasil by a unique process developed by Avco. Afer curing, the heatshield was machined to tolerance, the radar cross section reducing material applied and covered with a final layer of Avcoat 2. The aft fairing was specifically designed to reduce the radar cross section. (100)

While the Mark 4 and 5 tumbled at first during reentry and thus provided a large radar return, the Mark 11 was spin stabilized so as to present a reduced radar return for as long as possible. The Mark 11 deployed from the third stage with only a slight increase in velocity so the third stage served almost like a radar beacon for Soviet ABM systems.

Virtually indistinguishable in outer appearance, the Mark 11 series were approximately 100 inches in height, with cylindrical section nineteen inches in diameter, a base diameter of thirty-two inches and all used the same ablative material. The Mark 11 was deployed on Minuteman IB. All four variants were deployed on Minute-man II. For the Mark 11A and 11B, Avco developed a retro rocket spacer that had ten small thrusters which fired in pairs to provide a random velocity to the third stage. Before firing the retro rocket thrusters, a tumbler motor fired perpendicular to the centerline of the third stage to impart a rotation rate. This combination randomized the third stage position relative to the reentry vehicle and thus reduced the problem of the third stage serving as a radar beacon. (101)

For the Mark 11C the retrorocket spacer was replaced with a chaff spacer which carried a number of Mark 1A chaff dispensers, each equipped with different level impulse thrusters. This was in response to the low frequency Soviet ABM radars. The chaff dispenser was connected to the Mark 11C via a lightweight spacer made of beryllium rather than aluminum as this configuration was up against a weight limit due to the chaff system and beryllium was thirty percent lighter than aluminum. After Mark 11C release, the chaff dispensers were fired up and down the range insensitive axis to generate a train of chaff clouds spaced far enough apart that the defensive systems would have to target each cloud. (102)

The Mark 11 research and development program included six flights on Atlas D missiles beginning on August 28, 1963 and ending February 12, 1964 with one successful flight, the failures were due to booster malfunctions. Minuteman IB flight tests began on December 7, 1962 and ended on December 8, 1967 after forty-one flights with six failures. The Mark 11C penetration aids capability was tested on the final six flights which began on April 28, 1967. (103) (Figure 25).

The weight of the Mark 11 was 200-250 pounds. The Mark 11 A, B and C were twenty-five percent heavier than the Mark 11. (104) The Mark 11 series reentry vehicles were deployed on Minuteman IB and Minuteman II from 1963 to 1973 (Minuteman IB) and 1995 (Minute-man II). (105)

Summary

There were three key technologies that needed to be developed for the Minuteman program to succeed: large diameter solid propellant motors, lightweight inertial guidance systems and lightweight reentry vehicles. The evolution of reentry vehicle design began with the need to quickly design and field a reentry vehicle system for a relatively large warhead using readily available materials. The result was the first generation heatsink concept used with the Air Force Thor and Navy Polaris A-1 and A-2 IRBMs.

The second generation reentry vehicle system, ablation, was demonstrated first by the Army in its development of the reentry vehicle for the Jupiter IRBM. The Air Force quickly saw the advantage of ablation technology which permitted the design of lighter, more streamlined and hence more accurate, reentry vehicles. The Mark 5 and Mark 11 reentry vehicles represented the culmination of the pyrolytic or charring method of ablation with their small size and greater accuracy compared to heatsink reentry vehicle systems.

NOTES

(1.) T.C Lin, Development of U.S. Air Force Intercontinental Ballistic Missile Weapon Systems, Journal of Spacecraft and Rockets, Vol. 40, No. 4, July August 2003, 503-506.

(2.) W. Johnson, The Rise and Fall of Early War Rockets, International Journal of Impact Engineering, 1994, Vol. 15, No.4, 365-383; Winters, F, The First Golden Age of Rocketry: Congreve and the Hale Rockets of the 19th Century, (Washington, D.C.): 1990, Smithsonian Institution Press 1 9, 215 224.

(3.) W.E. Greene, The Development of the SM-68 Titan: Vol. I, Narrative, 1962, AFSC Historical Publication Series 62-23-1, footnote page 111. In March 1957, Avco proposed changing the term nose cone to reentry vehicle. Initial response was negative but the term was eventually accepted and is used throughout this text to avoid confusion.

(4.) W. Dornberger, V-2, (New York: Viking Press), 1954, 222; Gregory P. Kennedy, Vengeance Weapon 2, (Washington, D.C.: Smithsonian Institution Press, 1984), 68-69.

(5.) D.D. Baals, and W.R. Corliss Wind Tunnels of NASA, (National Aeronautics and Space Administration, Scientific and Technical Information Office, 1981) SP-440, 55; J.R. Hanson, Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958, (National Aeronautics and Space Administration, Scientific and Technical Information Office, 1986) SP-4305,343.

(6.) C. Wagner, The Skin Temperature of Missiles Entering the Atmosphere at Hypersonic Speed, (U.S. Department of the Army, Ordnance Research and Development Division), Technical Report No. 60 October 1949, 1.

(7.) H.J. Allen and A.J. Eggers, Jr., A Study of the Motion and Aerodynamic Heating of Missiles Entering the Earths Atmosphere at High Supersonic Speeds, National Advisory Committee on Aeronautics, 1953, RMA53D28, 17, 25 27. These two authors published several updated versions of this seminal paper. It is not clear to this author what was updated, the conclusions in each of the sequential papers appear to remain the same; E.P. Hartman, Adventures in Research, A History of Ames Research Center 1940-1965, (National Aeronautics and Space Administration, Scientific and Technical Information Office, 1970), NASA Center History Series, SP-4302, 216 218.

(8.) H.J. Allen, Hypersonic Flight and the Re-Entry Problem, Journal of the Aeronautical Sciences, Vol. 25, No. 4, 223, Fig. 9.

(9.) T. von Karman, Aerodynamic Heating - The Temperature Barrier in Aeronautics, in Proceedings of the Symposium on High Temperature - A Tool for the Future; Berkeley, California, June 1956, (Menlo Park, California: Stanford Research Institute, 1956), 140 142; CRC Handbook of Chemistry and Physics, 52nd Edition, 1971 72, (Cleveland, Ohio: The Chemical Rubber Company,1971), D-142.

(10.) A.J. Eggers, C.F. Hansen and B.E. Cunningham, Stagnation-Point Heat Transfer to Blunt Shapes in Hypersonic Flight, Including Effects of Yaw, National Advisory Committee for Aeronautics, 1958, Technical Note 4229, 1 2.

(11.) J.V. Becker, Results of Recent Hypersonic and Unsteady Flow Research at the Langley Aeronautical Laboratory, Journal of Applied Physics, (July 1950), 620.

(12.) D. Schmidt, Hypersonic Atmospheric Flight, in Environmental Effects of Polymeric Materials, (New York: Interscience Publishers, 1968), 489.

(13.) P. Vielle, Sur Les Discontinuities Produites par la Brusque de Gas Comprises , 1899, Comptes Rendus 129,(1899), 1228.

(14.) W. Payman and C.F. Shepard, Explosion Waves and Shock Waves. VI. The Disturbances Produced by Bursting Diaphragms with Compressed Air, 1946, Proceedings of the Royal Science Academy, Vol 186, (1946), 293.

(15.) B.D. Henshall,, On Some Aspects of the Use of Shock Tubes in Aerodynamic Research, (Aeronautical Research Council, Great Britian, 1957) Reports and Memoranda, No. 3044, 6; H.J. Davis and H.D. Curchack, Shock Tube Techniques and Instrumentation, (U.S. Army Material Command, Harry Diamond Laboratories, March 1969) TR-1429, 1 16. Available as Defense Technology Information Center AD692295.

(16.) Hartman, Adventures in Research, 241; J.A. Shortal, A New Dimension: Wallops Island Flight Test Range, the First Fifteen Years, National Aeronautics and Space Administration, Scientific and Technical Information Office, 1978) NASA Reference Publication 1028, 433.

(17.) G.H. Stine, ICBM: The Making of the Weapon That Changed the World, (New York: Orion Books, 1991), 187.

(18.) H.F. Swift, Light-Gas Gun Technology: A Historical Perspective, in High Pressure Shock Compressions of Solids, Vol. VIII, (Berlin Heidelberg, Germany, 2005), 1 6.

(19.) A.J. Eggers A Method for Simulating Atmospheric Entry of Long Range Ballistic Missiles, National Advisory Committee for Aeronautics, Ames Aeronautical Laboratory, 1958, TR-1378, 1009 1012.

(20.) Hartman, Adventures in Research, pp. 238, 264.

(21.) Ibid., 339.

(22.) JA. Fay and F.R. Ridell, Theory of Stagnation Point Heat Transfer in Dissociated Air, , Journal of Aeronautical Sciences, Vol 25, No 2, (February 1958), 73-85.

(23.) P.H. Rose and W. I. Stark, Stagnation Point Heat-Transfer Measurements in Dissociated Air, Journal of Aeronautical Sciences, Vol 25, No 2, (February 1958), 86-97.

(24.) L. Lees, Laminar Heat Transfer Over Blunt-Nosed Bodies at Hypersonic Speeds, Jet Propulsion, (April 1956), 259-269.

(25.) W.R. Lucas and J.E. Kingsbury, A Brief Review of the ABMA Ablation Material Program, (Redstone Arsenal,AL: Army Ballistic Missile Agency,12 May 1960) ABMA DSN-TM-7-60 1, 3.

(26.) Von Karman, Aerodynamic Heating, 142.

(27.) D.L. Schmidt, Ablative Plastics for Reentry Thermal Environments, (United States Air Force, Air Research and Development Command, Wright Air Development Division,1961), Rep 60-862, 4 5.

(28.) R.A. Fuhrman, The Fleet Ballistic Missile System: Polaris to Trident, Journal of Spacecraft, Vol 15, No. 5, (Oct-Sept, 1978), 270 271.

(29.) JR. Stoney and A.J. Swanson, Heat Transfer Measured on a Flat-Faced Cylinder in Free Flight at Mach Numbers Up to 13.9, NACA Research Memorandum L57ER13, 17 June 957, 1 2.

(30.) H.S. Carter and W.E. Bressette, Heat Transfer and Pressure Distribution on Six Blunt Noses at a Mach Number of 2, NACA Research Memorandum L57C18, 18 April 1957, 10.

(31.) J.R. Hall and B.J. Garland, A Feasibility Study of the Flare-Cylinder Configuration as a Reentry Body Shape for an Intermediate Range Missile, NACA Research Memorandum L58C21, 28 May 1958, 11.

(32.) Shortal, A New Dimension, 556-561; J.P. McManus, A History of the FBM System, (Lockheed Missiles and Space Company, Inc., 1989), A-11.

(33.) G.A. Hoffman, Materials for Space Flight, (Santa Monica, CA: RAND Corporation, P-1420, 1 July 1958), 1-26.

(34.) Rear Admiral Robert Wertheim, personal interview with author, May 2015; Fuhrman, The Fleet Ballistic Missile System, 272.

(35.) R.Smelt, Lockheed X-17 Rocket Test Vehicle and Its Applications, American Rocket Society Vol 29, No. 8, (1959), 565-567.

(36.) Shortal, A New Dimension, 377, 441 444.

(37.) Re-entry Research: The Lockheed X-17, Flight, 6 February 1959, 181; R.F. Piper, The Development of the SM-80 Minuteman, (U.S. Air Force Historical Office, Deputy Commander for Aerospace Systems, April, 1962), Memorandum for Secretary, Ordnance Technical Committee, Office of the Chief of Ordnance, Department of the Army, Subj: JATOs, T64 and T65 Establishment of DO A Project 517-1-032 (TU2-2032), 6 April 1955, Vol. II Supporting Documents, Document 1(1962), 2 4. The propellant weights differ slightly from other sources but this appears to be the original contract information.

(38.) E.S. Sutton, From Polymers to Propellants to Rockets: A History of Thiokol, in 35th AIAA/ASME/SAEA/ASEE Joint Propulsion Conference and Exhibit, (Los Angeles, California: American Institute of Aeronautics and Astronautics, June 20-24, 1999), Paper 99-2929, 13.

(39.) R.W.Roy and R.A Foster, Final Report: Re-Entry Test Vehicle X-17, 10 May 1957", History Air Force Missile Test Center 1 July - 31 December 1957, Vol TV Supporting Documents Appendix F, Air Force Historical Research Agency, IRIS NO. 484479, Reel 15345, K241.01, Vol. 4, 1 6, 30; T.A. Heppenheimer, Facing the Heat Barrier: A History of Hypersonics, (National Aeronautics and Space Administration, NASA History Office, 2007) NASA History Series, SP-2007-4232, 44.

(40.) Colonel Bogart Staff Study on R W Role, Air Force Historical Research Agency, IRIS 1040339, Reel 35267, 168.7171-191, Report No. 20, frame 211, no pagination; Recoverable Data Capsule Designed for Thor and Atlas Test Missiles, Aviation Week, 24 November 1958, 71 72.

(41.) The Scientific Advisory Committee's First Report on Ballistic Missiles to the Secretary of Defense, 11 February 1956, Digital National Security Archive, (httpyAvww.proquest.com/productsservices/databases/dnsa. html), NH00553, 3; Monthly Report on Progress of ICBM and IRBM Programs (Report No. 16), 31 March 1957, U.S. Declassified Documents Online (http://www.gale.com/c/us-declassified documents online, CK234914 9174, 2 July 2016), 6.

(42.) M. Morton, Progress in Reentry Recovery Vehicle Development (Philadelphia, PA: General Electric, Missile and Space Vehicle Division, 1961), 6. The author is indebted to Donald L. Schmidt for this document.

(43.) J.R. Stalder, The Useful Heat Capacity of Several Materials for Ballistic Nose Cone Construction, NACATN-4141, (November 1957), 1-8.

(44.) P. Fote, personnel interviews with author, July 2015, May 2017. Fote is Chief Engineer Missile Systems, Textron Systems Corporation. Fote was a reentry vehicle design engineer for Avco Corporation for the Mark 4, 5, and 11; Two Approaches Used in First Production Nose Cones, Aviation Week, 12 May 1958, 65.

(45.) K.C. Speegle, Pressure Measurement on an Afterbody Behind Various Blunt Nose Shapes at a Mach Number of 2, NACA RM L5714, 27 December 1957, 5.

(46.) Morton, Progress, 7; Ballistic Missile Program Progress Report, 15 October 1956, Eisenhower Library, Papers as Pres. of the US, 1953 61 (Ann Whitman File), Folder: Guided Missiles (5), Box 16, v.

(47.) Morton, Progress, 7; Holaday, W.M., Monthly Report on Progress of ICBM and IRBM Programs, Report 23, 31 October 1957, U.S. Declassified Documents Online (http://www.gale.com/c/us-declassified-documents-online, CK2349205928, 27 November 1957), 4.

(48.) D.J. Stewart, Nose Cones: The Case for Heat Sink, Missiles and Rockets, 8 June 1959, 16-17.

(49.) Smithsonian News Release, 14 May 1959; Recoverable Data Capsule Designed for Thor and Atlas Test Missiles, Aviation Week, 24 November 1958, 71 72.

(50.) Progress of ICBM and IRBM Programs, April, May, June 1960, United States Air Force, Office of the Director of Defense Research and Engineering, Digital National Security Archive, ttp://www.products-services/databases/dnsa.html), NH00710, 17; F.X. Ruggerio, Missileers Heritage (Air University, Maxwell AFB, 1981), Report Number 2065-81, 34-69.

(51.) R.H. Goddard, Report Concerning Further Developments, Smithsonian Institution Archive, March 1920, 4.

(52.) P. Juneau, Composite Materials, Ablative, Encyclopedia of Chemical Technology, 3rd Edition, (1978), Vol. I, 10-26. The author thanks Juneau for providing this reference, and for personal correspondence, February 1997.

(53.) E. Stuhlinger, Army Activities in Space, IRE Transactions in Military Electronics, Volume MIL, Issue 2, (April-July 1960), 65.

(54.) U.S. Congress, House of Representatives, Organization and Management of Missile Programs, Eleventh Report by the Committee on Government Operations: Hearings Before a Subcommittee of the Committee on Government Operations, 86th Congress, 1st session, 2 September 1959, 108.

(55.) E.J. Redman and L. Pasuk,, Pressure Distribution on an ABMA Jupiter Nose Cone (13.3 Degrees Semi-Vertex Angle) at Nominal Mach Numbers 5, 6, 7, and 8, NAVORD Report 4486, U.S. Naval Ordnance Laboratory Aeroballistic Research Report 381, 27 September 1957, 6. Available from Defense Technology Information Center, AD0158516; H.C. Dubose and R.C Bauer, Transonic Dynamic Stability Tests of a Half-Scale Model of the Jupiter W-14 Re-Entry Configuration, U.S. Air Force, Arnold Engineering Development Center,, AEDC TN-58-1, (February 1958), 3. Available from Defense Technology Information Center AD778200.

(56.) J.M. Grimwood and F. Strowd, History of the Jupiter Missile System, U.S. Army Ordnance Missile Command,(1962), Appendix 8, 8-1. Available from Defense Technology Information Center, ADA434109; Explorers in Orbit, Army Ballistic Missile Agency, 10 November 1958, National Archives and Record Administration, Record Group 156, Stack Area 290, Row 902, Compartment 9, Shelf 1, Box 10, 38-42 and 112-119; Lucas and Kingsbury, A Brief Review of the ABMA Ablation Materials Program, 6.

(57.) M.C. Cleary, Army Ballistic Missile Programs at Cape Canaveral 1953-1988, (45th Space Wing History Office,Patrick Air Force Base, Florida, 2006), 52; Grimwood and Strowd,, History of the Jupiter Missile System, Appendix 8, 8-1.

(58.) Ibid., Appendix 9, 9-2, 9-4.

(59.) W. von Braun, Letter to D.L. Schimdt, Technical Manager for Ablative Materials, Air Force Materials Laboratory, Wright-Patterson AFB, Ohio, 14 September 1969, National Archives and Research Administration, SE, RG 255, 01-0002, Box 24, 2; Grimwood and Strowd, History of the Jupiter Missile System, 65. Cutting off the tip of a right circular cone generates a frustum; J.C. Brassell, Jupiter: Development Aspects and Deployment, Mobile Air Material Area, Brookley AFB, Historical Office, Office of Information,1962, Air Force Historical Research Agency, IRIS 474833, Reel 14643, K205.0504-2, Vol 1, 30.

(60.) C. Hansen, The Swords of Armageddon: U.S. Nuclear Weapons Development Since 1945, Volume VII, (Sunnyvale, CA: Chuckelea Publications 1995), 323; M. Yaffee Two Approaches Used in First Production Nose Cones, 12 May 1958, Aviation Week, 61.

(61.) J. Neufeld, The Development of Ballistic Missiles in the United States Air Force 1945-1960, (Office of Air Force History 1990), Appendix 1: The Tea Pot Committee Report, 259-260; U.S. Congress, House of Representatives, Organization and Management of Missile Programs: Hearings Before a Subcommittee of the Committee on Government Operations, 3 March 1959, 383.

(62.) Monthly Report on Progress of ICBM and IRBM Programs, Report 21, 31 August 1957, Department of Defense, Office of the Director of Defense Research and Engineering, U.S. Declassified Documents Online (U.S. Declassified Documents Online tp://www.gale.com/c/us-declassified-documents-online,CK2349145648), 2.

(63.) W.E. Green, The Development of the SM-68 Titan, (History Office, Deputy Commander for Aerospace Systems, Air Force Systems Command, August 1962), Vol II. Supporting Documents: Letter, Brig Gen O J Ritland, V Cmdr, AFBMD, to Cmdr, ARDC,, Subject: Reorientation of Nose Cone Program, 22 Aug 1958, Air Force Historical Research Agency, IRIS 897235, Document 57, 1-2.

(64.) A.V. Levy, Evaluation of Reinforced Plastic Material in High Speed Guided Missile and Power Plant Application, Plastics World, Vol. 14, (March 1956), 10-11.

(65.) I.J. Gruntfest and L.H. Schenker, Behavior of Materials at Very High Temperatures,, Industrial and Engineering Chemistry, Vol. 50, (October 1958), 75A-76A; G.W. Sutton, Ablation of Reinforced Plastics in Supersonic Flow,, Journal of Aero/Space Sciences, (May 1960), 378. Refrasil is approximately 90 percent silicone oxide, made by leaching out the lower flux oxides of E glass fibers found in fiberglass.

(66.) M.C. Adams and E. Scala, The Interaction of High Temperature Air with Materials During Reentry, in: Proceedings of an International Symposium, on High Temperature Technology: Asilomar Conference Grounds, California, October 6 9, 1959, (New York: McGraw-Hill, 1960) 54 60; M. Yaffee, Ablation Wins Missile Performance Gain, Aviation Week, (18 July 1960), 57; Flight Summary Report Series D Atlas Missiles, (San Diego, CA: General Dynamics/Astronautics, 21 June 1961), 8-36. Available from Defense Technical Information Center, AD0833337.

(67.) U.S. Air Force, Study on the Future of AFBMD/MCO/SACMIKE/STL (R-W) Co., Air Force Historical Research Agency, IRIS # 01040315,168.7171-167, reel 35264, n.d., 74-79.

(68.) J.W. Powell, Thor-Able and Atlas Able, Journal of the British Interplanetary Society, Vol 37, (1984), 219 220, 222; Thor Able Launched in Nose Cone Test, Aviation Week, 28 July 1958, 27; D.L. Schmidt, Ablative Plastics for Thermal Protection, , Wright Air Development Division,U.S. Air Force, WADD Report 60-862, August 1968), 58. This report was provided to the author by Don Schmidt in June 1999; M. Morton, Progress, 11. The author thanks Don Schmidt for a copy of this paper.

(69.) W.M. Arms, Thor, the Workhorse of Space A Narrative History, (Huntington Beach, CA: McDonnell Douglas Astronautics Company, (31 July 1972), 4-1. The author thanks Joel Powell for a copy of this report.

(70.) Powell, Thor-Able, Atlas-Able, 220; Heppenheimer, Facing the Heat Barrier: A History of Hypersonics, 47; I. Clausen and E.A. Miller, Intelligence Revolution 1960: Retrieving the Corona Imagery that Helped Win the Cold War, (Washington, D.C.: U.S. Department of Defense, Center for the Study of National Reconnaissance, 26 April 2012), 46.

(71.) The author thanks Craig Brunetti for the Avco RVX-1 measurements.

(72.) Ablative Nose Cone Shows Reentry Effect, Aviation Week, 25 May 1959, 30.

(73.) Arms, Thor: The Workhorse of Space, A Narrative History, 4-4, 4-6; M. Morton, Thor-Able and Atlas Reentry Recovery Programs, (Philadelphia, PA: General Electric Defense Electronics Division, Missile and Space Vehicle Department, no date.), 1-7. The author is indebted to Don Schmidt for this document; Schmidt, Ablative Plastics for Reentry Thermal Protection, 58 59.

(74.) Flight Summary Report Series D Atlas Missiles, 2-1.

(75.) W.T. Barry, personal interview and correspondence with author, February 1997. Barry was a materials scientist consultant at the General Electric Space Sciences Laboratory in Philadelphia during the development of the Titan II Mark 6 reentry vehicle. The Series 100 plastic ablation process was patented by Barry, U.S. Patent 3,177,175 (1965).

(76.) While not explicitly described as such, the available pictures of the recovered vehicle appear to show material similar to the RVX-2A reentry vehicle which did have these materials.

(77.) Morton, Thor-Able and Atlas Reentry Recovery Programs, 22-23.

(78.) Barry, personal interview by author, February 1997.

(79.) Ibid..

(80.) Flight Summary Report, 8-38.

(81.) Ibid., 8-5 to 8-11.

(82.) Ibid., 8-2.

(83.) F.X Ruggerio, Missileers Heritage, 34-69.

(84.) Progress of ICBM and IRBM Programs, April, May, June 1960, Department of Defense, Office of the Director of Defense Research and Engineering, Declassified Documents Online (U.S. Declassified Documents Online tp://www.gale.com/c/us-declassified-documents-online, CK2349126998), 17.

(85.) W. Greene, The Development of the SM-68 Titan, 25.

(86.) Flight Test Summary, 8-32 to 8-37.

(87.) Progress of ICBM and IRBM Programs, April, May, June 1960, 17.

(88.) F.X. Ruggerio, Missileers Heritage, 34-69; J. Neufeld, Ballistic Missiles of the United State Air Force, 234, 276.

(89.) J. McDowell, personal interview by the author and extracts from unpublished History of Flight. The data can also be found at his website, http://www.planet4589.org/space/lvdb/index.html. McDowells database is considered the gold standard of civilian missile launch records.

(90.) Memos, May 1958, Personal Collection of Gen. B.A. Schriever, AFHRA, IRIS # 01040258, K168-7171-110, Reel 35259, 1-2.

(91.) R.F. Piper, The Development of the SM-80 Minuteman, Vol 2, Supporting Documents, Proposed Preliminary Operational Concept for Minuteman, 8 April 1958, Document 76, 6.

(92.) Ibid., Letter, Col H.L. Evans, Asst D/Cmdr, Space Sys, to Col C.H. Terhune D/Cmdr, Ballistic Missiles, 28 May 1958, Subject: Minuteman Warheads and Reentry Vehicles, Air Force Historical Research Agency, IRIS 897233, K243-012-5, Document 80, 1-2.

(93.) Personal interview by the author with Secretary of the Air Force Thomas Reed, May 2015. The companies were listed in his personal diary entry for Monday, 23 June 1958.

(94.) Space and Missile Systems Organization: A Chronology, 1954-1979, U.S. Air Force, Space Division/Chief of Staff History Office, October 1979), 57.

(95.) R.F. Piper, The Development of the SM-80 Minuteman, Vol 2, Supporting Documents, Cancellation of the Light Minuteman Reentry Vehicle, 9 December 1958, AFHRA IRIS#, 897233, K243-012-5 Document 94, 1.

(96.) Fote, personal interview by author, May 2017; Piper, The Development of the SM-80 Minuteman, Vol 1, 190 195.

(97.) McDowell, unpublished.

(98.) Operational Test Report: ELM BRANCH, Minuteman IA, LGM-30A, FTM 631, 1st Strategic Air Division, AFHRA, IRIS 918812, reel 26122, 2; J.C. Hopkins, The Development of the Strategic Air Command 1946 1981: A Chronological History, (Office of the Historian, Headquarters Strategic Air Command, 1 July 1982), 112, 157.

(99.) R.F. Piper, The Development of the SM-80 Minuteman, Vol I, 190 196.

(100.) Fote; M.L. Yaffee, Mark 11A Hardened Against Air Bursts, Aviation Week and Space Technology, 24 August 1964, 50 60.

(101.) Fote; M.L. Yaffee, Mark 11A Hardened Against Air Burst, 50-59.

(102.) Fote.

(103.) McDowell.

(104.) Operation Test Report: NICKED BLADE, Minuteman Missile IB, LGM-30B, FTM 1101, 1st Strategic Aerospace Division, AFHRA, IRIS 918790, reel 26121, 4; Fote.

(105.) Air Force News Service, http://www.af.mil/NewsATag/51024/air-force-news-service; Unites States Air Force Supporting Studies, A History of Strategic Arms Competition 1945 1972, June 1976, 362; J.C. Hopkins, The Development of the Strategic Air Command 1946 1981: A Chronological History, 191.

David K. Stumpf, Ph.D., is retired plant biochemist living with his wife, Susan, in Tucson, Arizona. He has written two nuclear weapon histories, Regulus the Forgotten Weapon, a history of the Navy's Regulus I and II cruise missiles and Titan II: A History of a Cold War Missile System, as well as contributing to the Air Force Missileers history. He is currently working on a, comprehensive history of the Minuteman ICBM program endorsed by the Office of the Secretary of the Air Force. He volunteered at the Titan Missile Museum, Sahuarita, Arizona, as museum historian and as a tour guide for 15 years. He was instrumental in the effort to gain National Historic Landmark status for the museum.
Table 1. Air Force Reentry Vehicle Designators Through Minuteman 11.

Mark 1           Atlas D, Thor         General Electric
                                       (development, not flown)
Mark 2           Atlas D, Thor         General Electric
Mark 3           Atlas D               General Electric
Mark 4           Atlas E, F. Titan I   Avco
Mark 5           Minuteman IA          Avco
Mark 6           Titan II              General Electric
Mark 7           Skybolt               General Electric (cancelled)
Mark 8, 9, 10    not assigned
Mark 11, 11 A,   Minuteman IB,         Avco
11B, 11C         Minuteman II

Miller, B., Studies of Penetration Aids Broadens, 20 January 1964,
Aviation Week and Space Technology, 79.
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Author:Stumpf, David K.
Publication:Air Power History
Geographic Code:1USA
Date:Sep 22, 2017
Words:17020
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