Showing papers on "Rocket published in 1972"

Optimal Control of Ordinary Differential Equations

Abstract: This chapter discusses the optimal control of ordinary differential equations. The formulation of the standard problem; existence of minimizing relaxed and approximate solutions; necessary conditions for a minimum; contingent equations and equivalent control functions; unbounded contingent sets and compactified parametric problems; variable initial conditions; and free time, infinite time, staging, and advance-delay differential problems are the major topics covered in the chapter. Most present-day space vehicles have multiple stages, that is, rocket engines mounted on top of one another with the payload on top of the highest stage. The lowest stage is activated first, and when its fuel is used up, it is jettisoned. When needed, the thrust is provided by the second stage and then the following ones. The control of a k -stage rocket thus involves k additional parameters, that is, the times when each of the rocket stages is discarded.

In-Flight Alignment and Calibration of Inertial Measurement Units - Part I: General Formulation

Abstract: This is the first part of a two-part paper which summarizes work pursued by the author in 1966 [1]. The paper describes the application of minimum-variance estimation techniques for in-flight alignment and calibration of an inertial measurement unit (IMU) relative to another IMU and/or some other reference. The first part formulates the problem, and the second part [2] reports numerical results and analyses. The approach taken is to cast the problem into the framework of Kalman-Bucy estimation theory, where velocity and position differences between the two IMU's are used as observations and the IMU parameters of interest become part of the state vector. Instrument quantization and computer roundoff errors are considered as measurement noise, and environmental induced random accelerations are considered as state noise. Typical applications of the technique presented might include the alignment and calibration of IMU's on aircraft carriers, the initialization of rockets or rocket airplanes which are launched from the wing of a mother ship, the alignment and calibration of IMU's which are only used in the latter phases of rocket flight, and for the initialization/updating of SST guidance systems.

Digital timer fuze

Abstract: A digital timer fuze (DTF) in subcombination with a digital timer fuze sym. The central element in the fuze system is the aircraft fire-control computer including a fuze setter, which receives input signals from various sensors as required for rocket trajectory computations and, in addition, pilot commands. When the pilot arms the system and selects a firing mode, power is supplied to each fuze in the load. When the firing button (pickle switch) is depressed, timing commands are injected into each fuze in proper sequence and the rocket motors are initiated via circuits and initiators incorporated in the launcher pods. As each rocket moves forward under motor thrust, the lead to its fuze is separated. Physical interruption of this circuit initiates the "run" phase of the digital timer fuze. The digital timer fuze is electrically connected to the fuze setter. A signal from the fuze setter charges a power supply capacitor and causes a counter to count clock pulses for a given period of time and store the count. When the umbilical line connecting the fuze to the fuze setter is severed, the main counter counts down at a given rate. When all of the stored counts have left the counter, the counter gates an SCR which allows the power supply capacitor to actuate the detonator.

Powder Rocket Feasibility Evaluation

Abstract: : The objective of this evaluation was to investigate the feasibility of the powder rocket concept. A series of rocket combustor tests were conducted at 500 and 900 pound thrust levels using ammonium perchlorate/aluminum propellants. Eight 500 pound thrust fire tests were conducted to determine the effect of propellant variations on combustion. Particle size affected combustion stability with smoother combustion obtained with reduced particle size. Combustion oscillations of plus or minus five percent of mean chamber pressure at a frequency of 140 Hertz occurred using three micron (X-65) Al and twenty micron AP which was the smoothest fire test. However, partial plugging of the AP injector occurred which was attributed to the agglomeration tendency of the finer AP. A test with AP containing 0.2 weight percent silica additive, improved the AP flow character but increased combustion oscillation level to plus or minus thirty-eight percent of mean Pc.

Comparison of advanced propulsion concepts for deep space exploration.

Abstract: Equations and charts are presented that permit rapid estimation of propulsion-system performance requirements for some typical deep-space missions. A number of advanced propulsion concepts for which performance estimates are available are compared with respect to their capability for flyby, rendezvous, and roundtrip planetary missions. Based on these estimates, the gas-core nuclear fission rocket and the pulsed fusion rocket yield the fastest trip times to the near planets. For round trips to Jupiter and beyond, the controlled fusion rocket shows progressively superior capabilities. Several propulsion concepts based on use of impinging laser beams are found to be noncompetitive with the other advanced concepts for deep space missions. Requirements for attainment of interstellar distances within a human lifetime are found to be some orders of magnitude beyond the capabilities of any propulsion concepts for which performance estimates are now possible.

A search for spectral features in the submillimeter background radiation.

Abstract: Mountaintop observations were made at 1 per cent spectral resolution of atmospheric and sky emission in an attempt to determine the spectral features of submillimeter background radiation. No emission features were found that could be related to the diffuse isotropic flux reported from rocket and balloon experiments.

Rocket Measurements of Ionization Irregularities in the Equatorial Ionosphere at Thumba & Identification of Plasma Instabilities

Abstract: Ji'Langmuir probe system capable of high frequency response has been designed and developed at the Physietlt Res~ LabM'ftt6Ty,Ahmedabad, and used at Thumba to study ionization irregularities in the equatorial lower ionosphere in the scale size ranges 1-15 and 30-300 m. Studies have been conducted at different times of the day as well as during night. It is found that during evening and night-time periods, many of the obs~ed features of the irregularities can be explained in terms of a drift dissipative instability mechanism connected with background electron density gradients and the presence of electric fields in

Fast Computer Program for Nonequilibrium Rocket Plume Predictions

Abstract: : A fast computer program for predicting nonequilibrium rocket plume properties is described. The analytical model assumes parallel turbulent (or laminar) mixing between concentric chemically reacting streams and can also be used for studying chemical lasers and re-entry wakes. The effects on radar attenuation are also presented. The equations for free shear layer mixing with nonequilibrium chemistry are solved. The stability problems inherent in fully explicit finite difference schemes are shown to be eliminated. Computer run times for typical afterburning rocket plume calculations are shown to be decreased over the original (fully explicit) AeroChemaxisymmetric mixing with nonequilibrium chemistry program. Both the analysis and computer program write- up are presented, including a sample calculation and a FORTRAN listing.

Dual thrust level monopropellant spacecraft propulsion system

Abstract: A flight auxiliary propulsion system for velocity trim, station keeping, momentum adjustment for a spacecraft comprising rocket or reaction motors, also designated thrusters, utilizing thermally decomposable monopropellants such as hydrazine and other derivatives, thereof hydrogen peroxide, and isopropyl nitrate. The thrusters are arranged in a distribution or manifold system so that one set of thrusters provides for relatively large thrusts of force in the order of 1 to 5 pounds and another set of thrusters develop low thrusts in the millipound range. The large thrusts are developed by the catalytic decomposition of the monopropellant into a thrust chamber and through a throat and expansion nozzle to the ambient externally of the spacecraft. The low level thrusts are developed by heating catalytically or thermally decomposed monopropellant by electrical heating elements more commonly known as resisto-jet elements. Dual thrust levels may also be achieved by a common motor with a controllable resisto-jet and variable throat-area control.

Pressurized Crack Behavior in Two-Dimensional Rocket Motor Geometries

Abstract: A method is presented for evaluating two-dimensional crack behavior in rocket motor geometries for pressure loadings in which the pressure is applied directly to the crack surfaces. The experimental requirements associated with pressurizing the crack necessitated the application of pressure over the two-dimensional plane surface of the specimen. Analytical solutions are developed which include the side pressure and relate the stress intensity factors to the classical unpressurized situation. Stress intensity factors for the complex cracked rocket motor geometries were evaluated using finite element computer techniques based on strain energy methods. Comparison between analytical predictions using elastic fracture mechanics and experimental observations of a brittle epoxy was quite good for the three different geometries tested. The work has application in fracture analysis of solid propellant rocket grains and in pressure vessels containing partial-thickness cracks which emanate from the inside.

To mars in 30 days by gas-core nuclear rocket.

Abstract: An important component for a concept involving a 60-day Mars mission is the gas-core nuclear-rocket engine. A gas-core reactor, however, has also other potential applications including MHD power generators, breeder reactors, and nuclear-powered lasers. The gas-core engine uses a fissioning uranium plasma to heat hydrogen and then expands it through a nozzle to convert the thermal energy into thrust. To obtain a higher specific impulse than the 825 sec of the solid-core nuclear-rocket engine, a gas core has to produce hotter hydrogen.

Further observations of the solar limb spectrum in the region 550–2000 Å

Abstract: Further observations of the ultraviolet spectrum (550–2000 A) of the solar limb and disc were obtained during a Skylark rocket flight on 5 August 1971. These observations have enabled several new spectral lines to be identified and classified.

Foldable kite and rocket launching means therefor

Abstract: A foldable kite adapted to be compactly arranged in a manner to have a small cross section for storage within the hollow interior of a minature, toy-like rocket having a rocket engine in the normally rearward end thereof and the opposite end of the rocket body having a conical nose removably mounted therein. The folded kite is positioned adjacent to rocket engine and the normally outer end of the rocket body has a compactly folded parachute stored therein and connected to the body of the rocket as well as to said nose. The rocket engine has a delay charge therein which, when the rocket has been projected to a predetermined height is fired automatically to project the nose and the parachute which is connected thereto, as well as the kite, and a tether cord extends from the kite down to a compact arrangement of such cord adapted to rapidly pay out as the rocket and kite ascend but permit a person to control the flight of the kite by said cord, while the rocket is recovered due to the parachute permitting its gradual decent without injury.

Foam structured rocket dispenser

Abstract: A rocket dispensing pod structure including a urethane foam enclosing a cluster of rocket launcher-tubes and wrapped in a reinforcing, outer skin cover composed of glass, graphite or other high modulus fibers covered with epoxy, polyester or other suitable resin.

Liquid fueled rocket engine of the main current type having separate control nozzles operated from turbine exhaust gases branched off from the rocket process

Abstract: A liquid fueled rocket engine of the so-called main current type in which combustion gases are generated in a precombustion chamber and directed in series to a fuel component pump drive turbine and to a main combustion chamber for generating thrust gases, includes one or more control nozzles or control nozzle groups which are connected to receive the exhaust gases from the turbine with the addition of small partial amounts of a propellant component.

Apparatus for staged combustion in air augmented rockets

Abstract: Apparatus for providing staged addition of air to the ramburner of an air gmented rocket for enhancing the ignition, combustion, and performance of the rocket. A shaped, strut-mounted, ablative-coated sleeve is mounted coaxially in the ramburner section of the rocket. This sleeve causes the diversion of a part of the inlet air and provides a favorable regime for ignition and combustion of the air augmented propellant. The diverted air is subsequently added to the combustion region through passageways in the sleeve. The shaped combustible ablative material of the sleeve provides additional thrust-producing fuel to the ramburner.

Performance Potential of Gas-Core and Fusion Rockets: A Mission Applications Survey

Abstract: This paper reports an evaluation of the performance potential of five nuclear rocket engines for four mission classes. These engines are: the regeneratively cooled gas-core nuclear rocket; the light bulb gas-core nuclear rocket; the space-radiator cooled gas-core nuclear rocket; the fusion rocket; and an advanced solid-core nuclear rocket which is included for comparison. The missions considered are: earth-to-orbit launch; near-earth space missions; close interplanetary missions; and distant interplanetary missions. For each of these missions, the capabilities of each rocket engine type are compared in terms of payload ratio for the earth launch mission or by the initial vehicle mass in earth orbit for space missions (a measure of initial cost). Other factors which might determine the engine choice are discussed. It is shown that a 60 day manned round trip to Mars is conceivable.-

Interactions between the Flow Field, Combustion, and Wave Motions in Rocket Motors

Abstract: : The stability of small amplitude oscillations in combustion chambers is analyzed for one- and three-dimensional problems. In addition to combustion and mass addition at the boundaries, residual combustion and the presence of particulate matter within the chamber are accounted for. The results for the one-dimensional problem introduce new contributions, to the balance of acoustic energy, associated essentially with boundary layer processes acting if there is a component of acoustical motion parallel to the surface. These are incorporated in the general three-dimensional problem, and are shown to have a significant influence on the predicted stability of motions in a rocket motor.

Rocket borne instrument to measure electric fields inside electrified clouds

Abstract: An apparatus for measuring the electric field in the atmosphere which includes a pair of sensors carried on a rocket for sensing the voltages in the atmosphere being measured. One of the sensors is an elongated probe having a fine point thereon, which causes a corona current to be produced as it passes through the electric field. An electric circuit is coupled between the probe and the other sensor and includes a high ohm resistor which linearizes the relationship between the corona current and the electric field being measured. A relaxation oscillator and transmitter are provided for generating and transmitting an electric signal having a frequency corresponding to the magnitude of the electric field.

Base drag reduction

Abstract: Means for reducing or eliminating the base drag of a rocket motor is provided by selected configurations of a drag reducing propellant. The drag reducing propellant is a slow burning propellant which is positioned within a rocket motor and burned to provide gases for relieving the partial vacuum experienced after completion of the boost or sustain propellant burning phase. At the end of the burning phase no gases are being ejected from the rocket nozzle. But the drag reducing propellant compensates for this condition by being a source for continued gas ejection whereby the partial vacuum is relieved and the base drag effect of a rocket motor is reduced or eliminated. The drag reducing propellant is tailored and configured for use in combination with a rocket motor having an end burning, start perforated, or cylindrical propellant grain. In an alternate design, the drag reducing propellant is employed in combination with a rocket motor where it is positioned between the rocket nozzle and the rocket shroud. When employed in this design, the drag reducing propellant is independent of the pressure inside the rocket motor case, a distinct advantage for certain combinations.