ALBERT

Description: The high strength-to-weight ratio of titanium alloys suggests their use for solid-propellant rocket-motor cases for high-performance orbiting or space-probe vehicles. The paper describes the fabrication of a 6-in.-diam., 0.025-in.-wall rocket-motor from the 6A1-4V titanium alloy. The rocket-motor case, used in the fourth stage of a successful JPL-NASA lunar-probe flight, was constructed using a design previously proven satisfactory for Type 410 stainless steel. The nature and scope of the problems peculiar to the use of the titanium alloy, which effected an average weight saving of 34%, are described.

Description: On 11 May 1959, 24 tests of the aerodynamic response of the McDonnell model Project Mercury capsule were conducted. The initial test demonstrated free-fall; a parachute was used in the remaining test. Several tests included the addition of baffles.

Description: An analysis is made of the oscillatory motion of vehicles which traverse arbitrarily prescribed trajectories through the atmosphere. Expressions for the oscillatory motion are derived as continuous functions of the properties of the trajectory. Results are applied to a study of the oscillatory behavior of re-entry vehicles which have decelerations that remain within limits of human tolerance. It is found that a deficiency of aerodynamic damping for such vehicles may have more serious consequences than it does for comparable ballistic missiles.

Description: Early in calendar year 1958 Space Technology Laboratories, Inc. (STL) (then Space Technology Laboratories, a division of the Ramo-Wooldridge Corp.) developed for the Air Force Ballistic Missile Division (AFBMD) an Advanced Re-entry Test Vehicle (ARTV) for the purpose of testing ballistic missile nose cones at the full range of 5500 nautical miles. The two-stage ARTV utilized the Thor ballistic missile and the second stage propulsion system developed for the Vanguard program. In late 1957 and early 1958, STL/AFBMD prepared studies of various missile combinations which could be utilized for space testing. The Thor, in combination with the Vanguard second and third stages, was one of the vehicles considered which offered a very early capability of placing a reasonable payload in a lunar orbit. These STL/AFBMD studies were presented to various appropriate groups including the Killian, Millikan, H. J . Stewart Committees; Headquarters, Air Research and Development Command, and ARDC Centers. Subsequently the Advanced Research Projects Agency (ARPA) contacted STL relative to the availability of hardware for an early lunar shot. By utilizing existing spares already purchased for the ARTV, and by making use of the ARTV contractors already in being, it appeared feasible to launch by the third quarter of calendar year 1958 a payload which would be captured by the moon's gravitational force. On 27 March 1958, ARPA directed STL to proceed with a program of three lunar shots. As much as possible, these shots were to utilize existing ARTV spare hardware and impose no interference with the ballistic missile programs. In September this program was transferred to the direction of the National Aeronautics and Space Administration (NASA). On 17 August 1958 the first launching of the Able-1 vehicle was attempted, but the flight was terminated by a propulsion failure of the first stage. Subsequent launchings were attempted on 13 October and 8 November 1958. Of these launchirigs the October attempt was the most successful. Although the payload did not reach the vicinity of the moon, a maximum altitude of 71,700 was attained, and useful scientific data was obtained from the instrumentation.

Description: A brief theoretical study has been made for the purpose for estimating and comparing the weight of three different types of controls that can be used to change the attitude of a satellite. The three types of controls are jet reaction, inertia wheel, and a magnetic bar which interacts with the magnetic field of the earth. An idealized task which imposed severe requirements on the angular motion of the satellite was used as the basis for comparison. The results showed that a control for one axis can be devised which will weigh less than 1 percent of the total weight of the satellite. The inertia-wheel system offers weight-saving possibilities if a large number of cycles of operation are required, whereas the jet system would be preferred if a limited number of cycles are required. The magnetic-bar control requires such a large magnet that it is impractical for the example application but might be of value for supplying small trimming moments about certain axes.

Description: A concept for a manned satellite reentry from a near space orbit and a glide landing on a normal size airfield is presented. The reentry vehicle configuration suitable for this concept would employ a variable geometry feature in order that the reentry could be made at 90 deg. angle of attack and the landing could be made with a conventional glide approach. Calculated results for reentry at a flight-path angle of -1 deg. show that with an accuracy of 1 percent in the impulse of a retrorocket, the desired flight-path angle at reentry can be controlled within 0.02 deg. and the distance traveled to the reentry point, within 100 miles. The reentry point is arbitrarily defined as the point at which the satellite passes through an altitude of about 70 miles. Misalignment of the retrorocket by 10 deg. increased these errors by as much as 0.02 deg. and 500 miles. Intra-atmospheric trajectory calculations show that pure drag reentries starting with flight-path angles of -1 deg. or less produce a peak deceleration of 8g. Lift created by varying the angle of attack between 90 and 60 deg. is effective in decreasing the maximum deceleration and allows the range to the "recovery" point (where transition is made from reentry to gliding flight) to be increased by as much as 2,300 miles. A sideslip angle of 30 deg. allows lateral displacement of the flight path by as much as 60 deg. miles. Reaction controls would provide control-attitude alignment during the orbit phase. For the reentry phase this configuration should have low static longitudinal and roll stability in the 90 deg. angle-of-attack attitude. Control could be effected by leading-edge and trailing-edge flaps. Transition into the landing phase would be accomplished at an altitude of about 100,000 feet by unfolding the outer wing panels and pitching over to low angles of attack. Calculations indicate that glides can be made from the recovery point to airfields at ranges of from 150 to 200 miles, depending upon the orientation with respect to the original course.

Description: The three NASA/USAF lunar probes of August 17, October 13, and November 8, 1958 are described. Details of the program, the vehicles, the payloads, the firings, the tracking, and the results are presented. Principal result was the first experimental verification of a confined radiation zone of the type postulated by Van Allen and others.

Description: Variables for the reentry capsule water landing tests were flight path, vertical contact velocity, and contact attitude. The capsule weighed 1900 pounds with a center of gravity 16.8 inches above maximum diameter.

Description: Methods of controlling the trajectories of high-drag-low-lift vehicles entering the earth's atmosphere at angles of attack near 90 deg and at initial entry angles up to 3 deg are studied. The trajectories are calculated for vehicles whose angle of attack can be held constant at some specified value or can be perfectly controlled as a function of some measured quantity along the trajectory. The results might be applied in the design of automatic control systems or in the design of instruments which will give the human pilot sufficient information to control his trajectory properly during an atmospheric entry. Trajectory data are compared on the basis of the deceleration, range, angle of attack, and, in some cases, the rate of descent. The aerodynamic heat-transfer rate and skin temperature of a vehicle with a simple heat-sink type of structure are calculated for trajectories made with several types of control functions. For the range of entry angles considered, it is found that the angle of attack can be controlled to restrict the deceleration down to an arbitrarily chosen level of 3g. All the control functions tried are successful in reducing the maximum deceleration to the desired level. However, in order to avoid a tendency for the deceleration to reach an initial peak decrease, and then reach a second peak, some anticipation is required in the control function so that the change in angle of attack will lead the change in deceleration. When the angle of attack is controlled in the aforementioned manner, the maximum rate of aerodynamic heat transfer to the skin is reduced, the maximum skin temperature of the vehicle is virtually unaffected, and the total heat absorbed is slightly increased. The increase in total heat can be minimized, however, by maintaining the maximum desired deceleration for as much of the trajectory as possible. From an initial angle of attack of 90 deg, the angle-of-attack requirements necessary to maintain constant values of deceleration (1g to 4g) and constant values of rate of descent (450 to 1,130 ft/sec) as long as it is aerodynamically practical are calculated and are found to be moderate in both magnitude and rate. Entry trajectories made with these types of control are presented and discussed.

Description: Early in 1958, the Jet Propulsion Laboratory of the California Institute of Technology was requested to participate in a lunar-probe mission code-named Juno II which would place a 15-lb Instrumented payload (Pioneer IV) in the vicinity of the moon. The vehicle was to use the same high-speed upper-stage assembly as flown on the successful Jupiter-C configuration; however, the first-stage booster was to be a Jupiter rather than a Redstone. An analysis of the intended flight and payload configuration Indicated that the feasibility of accomplishing the mission was questionable and that additional performance would have to be obtained if the mission was to be feasible. Since the most efficient way of Increasing the performance of a staged vehicle is to increase the performance of the last stage, a study of possible ways of doing this was made.. Because of the time schedule placed on this effort It was decided to reduce the weight of the fourth-stage rocket-motor case by substituting the annealed 6Al--4V titanium alloy for the Type 410 stainless steel. Although this introduced an unfamiliar material, It reduced the changes in design and fabrication techniques. This particular titanium alloy was chosen on the basis of previous tests which proved the suitability of the alloy as a pressure-vessel material when used at an annealed yield strength of about 120, 000 psi. The titanium-case fourth stage of Juno U is shown with the payload and on the missile in Fig. 1; the stainless-steel motor cases used in the Jupiter-C vehicle are shown in Fig. 2. The fourth-stage motor case has a diameter of 6 in., a length of approximately 38 in. center dot and a nominal cylindrical wall thickness of 0.025 in. As shown in Fig. 1, the case serves as the structural support of the payload and is aligned to the upper stage assembly through an alignment ring. The nozzle is threaded into the end of the motor case, and is of the ceramic-coated steel design. Figure 3 shows a comparison of the components used to make the stainless steel and the 6A1--4V titanium alloy cases. The forward dome and aft fitting for the stainless steel assembly were fabricated from a combination of forged, spun and machined parts.. In order to facilitate the fabrication of the titanium alloy motor ) these components were machined from a large-diameter billet.

Description: Simplified expressions describing the transfer from a satellite bit to the point of atmospheric entry are derived. The expressions are limited to altitude changes that are small compared with the earth's radius, and velocity changes small compared with satellite velocity. They are further restricted to motion about a spherical, nonrotating earth. The transfer orbit resulting from the application of thrust in any direction at any point in an elliptic orbit is considered. Expressions for the errors in distance (miss distance) and entry angle due to an initial misalignment and magnitude error of the deflecting thrust are presented. The largest potential contributing factor towards a miss distance stems from the misalignment of the retrovelocity increment. If this velocity increment is pointed in direct opposition to the path, a 1 deg misalignment leads to a miss distance of 3.45 miles. However, it is shown that this error can be avoided by applying the velocity increment at an angle between 120 deg and 150 deg below the flight-path, direction. The guidance and accuracy requirements to establish a circular orbit, in addition to the corrections applied to to transform elliptic orbits into circular ones, are also discussed.

Description: The three NASA/USAF lunar probes of August 17, October 13, and November 8, 1958 are described. Details of the program, the vehicles, the payloads, the firings, the tracking, and the results are presented. Principal result was the first experimental verification of a confined radiation zone of the type postulated by Van Allen and others.