Missile stage separation indicator and stage initiator Patent
Piezoelectric means for missile stage separation indication and stage initiation
Mayo, J. W.↗
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Static Longitudinal and Lateral Stability Parameters of Three Flared-skirt Two-stage Missile Configurations at a Mach Number of 6.86
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Penland, Jim A↗
Performance analysis of SA-3 missile second stage
One SA-3 missile was disassembled. The constituents of the second stage were thoroughly investigated for geometrical details. The second stage slotted composite propellant grain was subjected to mechanical properties testing, physiochemical analyses, and burning rate measurements at different conditions. To determine the propellant performance parameters, the slotted composite propellant grain was machined into a set of small-size tubular grains. These grains were fired in a small size rocket motor with a set of interchangeable nozzles with different throat diameters. The firings were carried out at three different conditions. The data from test motor firings, physiochemical properties of the propellant, burning rate measurement results and geometrical details of the second stage motor, were used as input data in a computer program to compute the internal ballistic characteristics of the second stage.
Helmy, A. M.↗
Buckling analysis of a missile interstage
The NASTRAN program was used to find the buckling location and the degradation factor due to a specified temperature field for an interstage of a typical three stage missile. The model was made up of 6332 independent degrees of freedom. This large size required the use of an IBM 370/195 for evaluation of the eigenvalues. The effect of elevated temperature was to lower the critical buckling load by 3%.
Dudley, D. P.↗
SBASI: Actuated pyrotechnic time delay initiator
A precision pyrotechnic time delay initiator for missile staging was developed and tested. Incorporated in the assembly is a single bridgewire Apollo standard initiator (SBASI) for initiation, a through-bulkhead-initiator to provide isolation of the SBASI output from the delay, the pyrotechnic delay, and an output charge. An attempt was made to control both primary and secondary variables affecting functional performance of the delay initiator. Design and functional limit exploration was performed to establish tolerance levels on manufacturing and assembling operations. The test results demonstrate a 2% coefficient of variation at any one temperature and an overall 2.7% coefficient of variation throughout the temperature range of 30 to 120 F. Tests were conducted at simulated operational altitude from sea level to 200,000 feet.
Salter, S. J.↗
Dynamic characteristics of a two-stage variable-mass flexible missile with internal flow
A general formulation of the dynamical problems associated with powered flight of a two stage flexible, variable-mass missile with internal flow, discrete masses, and aerodynamic forces is presented. The formulation comprises six ordinary differential equations for the rigid body motion, 3n ordinary differential equations for the n discrete masses and three partial differential equations with the appropriate boundary conditions for the elastic motion. This set of equations is modified to represent a single stage flexible, variable-mass missile with internal flow and aerodynamic forces. The rigid-body motion consists then of three translations and three rotations, whereas the elastic motion is defined by one longitudinal and two flexural displacements, the latter about two orthogonal transverse axes. The differential equations are nonlinear and, in addition, they possess time-dependent coefficients due to the mass variation.
Meirovitch, L.↗
A variational method for optimal staging.
Optimal steering and staging of multistage boosters using extended variational method to include variable discontinuities at corners
MULTISTAGE ROCKET↗
A variational method for optimal staging.
Optimal steering and staging of multistage boosters using extended variational method to include variable discontinuities at corners
MULTISTAGE ROCKET↗
Control characteristics for wrap-around fins on cruise missiles configurations
This paper presents selected results of a panel loads study conducted as part of the final phase of an extensive investigation of an air-breathing missile concept employing wrap-around aerodynamic surfaces. Typical results for M = 2.36 are presented for the fin load results, plus a brief review of basic results of the previously reported tests. Vapor screen results are also discussed. The present results indicate that the fin load characteristics are nearly identical for planar and curved fins having the same projected planform and would permit the use of planar-surface predictions for supersonic speeds in the preliminary design stages of missiles employing wrap-around curved fins.
Sawyer, W. C.↗
Dynamic characteristics of a variable-mass flexible missile: Dynamics of a two-stage variable-mass flexible rocket
The dynamic characteristics of two-stage slender elastic body were investigated. The first stage, containing a solid-fuel rocket, possesses variable mass while the second stage, envisioned as a flexible case, contains packaged instruments of constant mass. The mathematical formulation was in terms of vector equations of motion transformed by a variational principle into sets of scalar differential equations in terms of generalized coordinates. Solutions to the complete equations were obtained numerically by means of finite difference techniques. The problem has been programmed in the FORTRAN 4 language and solved on an IBM 360/50 computer. Results for limited cases are presented showing the nature of the solutions.
Meirovitch, L.↗
A Limited Study of a Hypothetical Winged Anti-ICBM Point-Defense Missile
A preliminary investigation was conducted to determine whether a warhead stage of an antimissile missile could be placed within an arbitrary 2-nautical-mile-radius maneuver cylinder around an intercontinental-ballistic-missile (ICBM) flight path above an altitude of 140,000 feet, a horizontal range of 40 nautical miles, at a flight-path angle of approximately 20 deg, and within 50 seconds after take-off using only aerodynamic forces to turn the antimissile missile. The preliminary investigation indicated that an antimissile missile using aerodynamic forces for turning was capable of intercepting the ICBM for the stated conditions of this study although the turning must be completed below an altitude of approximately 70,000 feet to insure that the antimissile missile will be at the desired flight-path angle. Trim lift coefficients on the order of 2 to 3 and a maximum normal-acceleration force of from 25g to 35g were necessary to place the warhead stage in intercept position. The preliminary investigation indicated that for the two boosters investigated the booster having a burning time of 10 seconds gave greater range up the ICBM flight path than did the booster having a burning time of 15 seconds for the same trim lift coefficient and required the least trim lift coefficient for the same range.
Brown, Clarence A., Jr.↗
Application of Describing-Function Analysis to the Study of an On-Off Reaction-Control System
An analytical study was made of an automatic reaction-control system for the upper stages of a missile to determine limit cycle characteristics, corresponding duty cycles, and the effects of the various system parameters on these quantities. A nonlinear servo analysis (describing function) technique was used to obtain a mathematical representation of the nonlinear components of the system. The results obtained by this analysis are compared with the results obtained from an analog simulation including the jet reaction-control hardware. The good agreement between the results of the two studies tends to indicate the feasibility of using such an analysis technique in the early design phase of a reaction-control system, since by so doing, the parameters for good system performance can be determined quite readily.
Lineberry, Edgar C., Jr.↗
1958 NASA/USAF Space Probes (Able-1): Summary - Volume 1
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.
Source record↗
The Explorer Rocket Research Program
Since September of 1956, nine Jupiter-C missiles have been launched from the firing pad at Cape Canaveral. The first Jupiter-C firing tested the propulsion system, air frame, and guidance components of the missile, and the second and third firings tested a model of the Jupiter nose cone under realistic re-entry conditions. The remaining six Jupiter-C missiles were used as the launching vehicles for EXPLORER satellites I through VI (Fig. 1). Of the six satellite firings, EXPLORERs I, III, and IV achieved satisfactory orbits. The Jupiter-C missile was designed and developed as a joint program under the technical direction of the Jet Propulsion Laboratory and the Army Ballistic Missile Agency. The Jet Propulsion Laboratory developed the three high-speed stages, and the Army Ballistic Missile Agency handled the development, construction, and operation of the first-stage booster rocket and the guidance system. Many other organizations have contributed to the success of the EXPLORER satellite program, most notably the State University of Iowa, the Air Force Cambridge Research Center, and the satellite tracking teams of the Vanguard organization.
Robillard, G.↗
Worldwide Space Launch Vehicles and Their Mainstage Liquid Rocket Propulsion
Space launch vehicle begins with a basic propulsion stage, and serves as a missile or small launch vehicle; many are traceable to the 1945 German A-4. Increasing stage size, and increasingly energetic propulsion allows for heavier payloads and greater. Earth to Orbit lift capability. Liquid rocket propulsion began with use of storable (UDMH/N2O4) and evolved to high performing cryogenics (LOX/RP, and LOX/LH). Growth versions of SLV's rely on strap-on propulsive stages of either solid propellants or liquid propellants.
Rahman, Shamim A.↗
Missileborne Artificial Vision System (MAVIS)
Several years ago when INTEL and China Lake designed the ETANN chip, analog VLSI appeared to be the only way to do high density neural computing. In the last five years, however, digital parallel processing chips capable of performing neural computation functions have evolved to the point of rough equality with analog chips in system level computational density. The Naval Air Warfare Center, China Lake, has developed a real time, hardware and software system designed to implement and evaluate biologically inspired retinal and cortical models. The hardware is based on the Adaptive Solutions Inc. massively parallel CNAPS system COHO boards. Each COHO board is a standard size 6U VME card featuring 256 fixed point, RISC processors running at 20 MHz in a SIMD configuration. Each COHO board has a companion board built to support a real time VSB interface to an imaging seeker, a NTSC camera, and to other COHO boards. The system is designed to have multiple SIMD machines each performing different corticomorphic functions. The system level software has been developed which allows a high level description of corticomorphic structures to be translated into the native microcode of the CNAPS chips. Corticomorphic structures are those neural structures with a form similar to that of the retina, the lateral geniculate nucleus, or the visual cortex. This real time hardware system is designed to be shrunk into a volume compatible with air launched tactical missiles. Initial versions of the software and hardware have been completed and are in the early stages of integration with a missile seeker.
Andes, David K.↗
Analytical Investigation of an Adaptive Flight-Control System Using a Sinusoidal Test Signal
An analytical study was made of an adaptive flight-control system which measures vehicle response to small-amplitude control-surface deflections produced by a sinusoidal test signal. Changes in the response to this signal are related to environmental changes,, and the system is continuously altered to maintain this response equal to a preselected value. The system is suitable for use in high-performance aircraft and missiles and requires only the addition of a signal generator and a logic circuit consisting of a filter-rectifier network and a comparator-integrator network to a basic command-control system. Thus, it presents a relatively simple approach to the problem. The effects on system performance of variation in flight condition, system-gain level, test-signal frequency, and sensor location are included in the analysis. Longitudinal control of a high-performance research aircraft over flight conditions ranging from landing approach to a Mach number of 5.8 at an altitude of 150,000 feet, and longitudinal control of a four-stage solid-fuel missile including the first bending mode over the atmospheric portion of a launch trajectory constituted the basis for the analytical study. Results of an analog-computer study using time-varying coefficients are presented to compare the control obtained with the adaptive system with-that obtained with a fixed-gain system during the atmospheric portion of a missile launch trajectory. The system has demonstrated an ability to maintain satisfactory vehicle control-system stability over wide ranges of environmental change.
Harris, Jack E.↗
Conceptual study of hypersonic airbreathing missiles
The purpose of this paper is to report recent results of an in-house conceptual study to evaluate the performance potential and research needs of airbreathing hypersonic missiles. An alkylated-borane (noncryogenic) fueled, dual-mode, ramjet/scramjet propulsion system structured with a Rene 41 inlet and a carbon-carbon combustor was assumed along with a Lockalloy heat sink fuselage structure and beryllium wings and control surfaces. Performance for an air-launched baseline missile with a 961 pound staging weight containing a 100 pound payload indicated excellent long range cruise, moderate acceleration and high maneuverability potential. A sizing study indicates that Mach 6 cruise ranges of the order of 2500 nautical miles for payloads of 300 pounds can be achieved with moderate size missile carry weights (9000 lbs.). Aerodynamic heating analyses indicate that unprotected heat-sink structures with internal insulation are feasible for ranges of several hundred miles. For ranges of several thousands of miles a multiwall radiation shield (Inconel/titanium) was selected for protection of the internally insulated heat sink structure.
Hunt, J. L.↗