Showing papers on "Rocket published in 1974"

Accuracy of rocket measurements of lower ionosphere electron concentrations

Abstract: Measurements of electron concentration taken at the same time and at the same place in the lower ionosphere by independent instrumentation mounted on the same rocket are described. The technique utilizes Faraday rotation and differential absorption of radio waves propagating from the ground to the rocket at two different frequencies. Agreement near 90 km within 7%, 6%, 8%, and 3% is demonstrated by the four available cases of coincidence in time and altitude. Maximum dispersion at other altitudes is calculated from known random errors. Stronger variation of electron collision frequency with altitude than with season is indicated by 34 measurements between 75 and 100 km. Insensitivity of electron concentration determinations at 72 km to errors in extrapolated collision frequency models is demonstrated.

Electron Echo Experiment 1: Comparison of observed and theoretical motion of artificially injected electrons in the magnetosphere

Abstract: On August 13, 1970, a sounding rocket launched from Wallops Island, Virginia, was used to successfully inject a controlled beam of energetic electrons into the trapping region of the earth's magnetosphere. The Aerobee 350 rocket carried a high-voltage electron gun, electron detectors, receivers to measure the electric field of waves generated by the experiment, and other apparatus to a peak altitude of 350 km. During a 10-min flight, more than 3000 16-ms-long 70-mA pulses of 35- to 43-keV electrons were injected into the magnetosphere with pitch angles within 25° of the local 90° mirror angle. The electrons, trapped in the geomagnetic field, bounced between the mirror point near the rocket and the southern conjugate point (−66° latitude, −82° longitude) while they were drifting eastward across field lines on a drift shell of constant McIlwain L parameter (2.56) under the gradient and curvature forces. Since Wallops Island conjugate mirror altitudes are at least 300 km lower than the injection altitudes, all returning electrons were backscattered from the atmosphere in the southern hemisphere. The rocket was launched eastward in an attempt to intercept the returning electrons, and for a 90-s interval about L apogee, electrons were detected returning from one, two, and in one case, three complete bounces to the southern hemisphere. The analyzed electron data agree well with echo patterns predicted from theoretical particle motion in models of the geomagnetic field combined with Monte Carlo studies of energy loss and diffusion in the conjugate point atmosphere. Observed bounce times, drift displacements, and cross L diffusion were in substantial agreement with theory. However, the returning flux was lower than predicted, and discrepancies between theory and experiment were observed at large diffusion distances. This experiment, Electron Echo 1, verified the concept of using controlled particle beams as magnetospheric probes, showing that such beams are sufficiently stable to be intercepted after bounces to the conjugate point and back.

An investigation of wave-particle interactions and particle dynamics using electron beams injected from sounding rockets

Abstract: Electrons with energy up to 40 kV have been injected into semi-trapped orbits from sounding rockets at Wallops Island, Virginia, and at Fort Churchill, Manitoba, Canada By directing the rocket trajectory to have a horizontal component which in direction and speed matched the bounce displacement of the injected electrons, it was possible to detect conjugate echoes at Wallops and possibly at Churchill, and to study the distribution of the echoes in space, time and energy By combining observations of many echoes, a composite picture can be obtained of the beam patterns Atmospheric scattering at the conjugate point of Wallops Island has been extensively studied, and the process produces a scale width of 5-10 m, but with an echo intensity which is only 10% of theoretical estimates

Ramjet with integrated rocket boost motor

Abstract: A combination rocket motor ramjet engine having the rocket motor located in the ramjet inlet diffuser. All or part of the rocket portion of the apparatus remaining after rocket booster burnout are ejected by air pressure in the diffuser inlet to permit the diffuser to supply ram air for the ramjet engine.

Mixed mode rocket engine

Abstract: Four embodiments of high-thrust rocket engines, each having separate propellant delivery apparatus for each propellant to be delivered to the engine thrust chamber, are illustrated herein. One engine is designed to provide a high thrust using only one fuel and one oxidizer. The other three engine embodiments are all designed to use different propellants having different densities during different portions of a rocket flight. Each of these additional engines includes the basic structure of the first. The first engine is thus not only an effective engine, but also a building block that can be easily modified to provide an engine capable of operating effectively with different propellants.

Low‐altitude acceleration of auroral electrons during breakup observed by a mother‐daughter rocket

Abstract: By the use of a mother-daughter rocket combination and ground-based observations with television, time and space variations are resolved in particle measurements in breakup aurora. The spectral variations measured during a temporal variation in the aurora can be explained by a nearly uniform acceleration of all the electrons such as would be caused by an electric potential drop along the magnetic field lines. Many other explanations can be eliminated.

Light-weight rocket deployable gas generator

Abstract: A compact, light weight gas generator employing the reaction of a solid grain mixture initiated by a primer Gases evolved are cooled and filtered through a series of filter disks

Two stage rocket with pressure responsive means for frictionally engaging second stage

Abstract: A toy rocket has a motor receptacle designed to receive a rocket motor of the type fueled by a self-pressurizing liquid propellant. The rocket motor during the fueling operation has its nozzle seated in a launcher through which liquid propellant flows into the propellant cavity of the rocket motor. When the propellant cavity of the rocket motor is pressurized, a member responsive to propellant pressure in the rocket motor engages a portion of the walls of the motor receptacle in the rocket body. When the pressure responsive member releases the receptacle walls, relative motion is provided between at least a portion of the rocket body and the rocket motor. For example, in one embodiment of the toy rocket, a second stage portion of the rocket is spring-biased away from a first stage portion of the rocket body. The major section of the rocket motor is fixed to the first stage portion. The pressure responsive member on the rocket motor engages the second stage portion while the rocket motor contains propellant to prevent separation of the first and second stages. When the propellant is expended, the pressure responsive member releases the second stage, allowing separation of the second stage from the first stage. A parachute is then deployed from the second stage to brake the descent of the rocket.

Solar x‐ray photography with multiplex pin‐hole camera

Abstract: A scatter‐hole x‐ray camera has been designed and flown in a rocket to measure solar x radiation. Its distinguishing feature is that many single pin‐hole images are allowed to overlap—a multiplexing approach that saves space, gives practical signal‐to‐fog ratio, and requires special object reconstruction techniques. It was possible to reconstruct the appearance of the dominant source of solar emission on the day of rocket flight with better than one arc minute resolution. The coronal x‐ray region had the same general shape and intensity distribution as the associated calcium K‐line region

Deployment of conductors into the atmosphere

Abstract: A system for rapidly deploying long, electrically conducting filaments of metallized Mylar in the air at selected altitudes by winding the Mylar in tight rolls and loading the rolls into small rocket pods in the nose of a folding fin aircraft rocket. The rolls are pyrotechnically ejected by causing the pod to fly apart and allow the tape rolls to enter the air stream and unwind.

Near‐simultaneous measurement of low‐energy electrons by sounding rocket and satellite

Abstract: An auroral sounding rocket was launched from Fort Churchill at 0453 UT on April 24, 1968. At the time of the rocket flight, the Vela 3B satellite was passing through the magnetotail plasma sheet at very nearly the same magnetic local time. Electron spectra obtained from experiments on these vehicles are well described by central Maxwell-Boltzmann populations, with indications of a high-energy tail; the satellite spectra also suggest the presence of a low-energy tail. The higher-temperature spectra measured on the rocket may be derived from the satellite spectra by adiabatic compression.

Model rocket-glider

Abstract: A solid fuel-powered model rocket having a main body tube and at least one auxiliary body tube. A solid fuel motor is mounted in the after end of one of the tubes and a slidable plug is disposed in the after end of the other tube, the plug being provided with a latch for engaging a wing elevator and maintaining it parallel to the horizontal plane of the wing against the biasing of a resilient band while the craft is in vertical flight. Following burn-out of the propellant charge, the gases generated upon ignition of a blow-out charge in the fuel cartridge cause rearward displacement of the slidable plug, disengagement of the latch and deployment of the elevator to a predetermined up-angel relative to the top horizontal plane of the wing, thereby directing the craft into a generally horizontal glide attitude.

Viscoelastic rocket grain fracture analysis

Abstract: A viscoelastic fracture analysis has been developed for rocket grain fracture predictions. The fracture analysis uses a stress intensity factor technique to predict crack velocity histories under thermal and pressurization loading conditions. The theory is compared with two-dimensional pressurized tests of two typical rocket motor geometries using the viscoelastic material, Solithane 113.

Rocket nozzle construction and surfaces impervious to hot, high velocity gases

Abstract: A thrust nozzle for a rocket has an improved exit cone made of a strong, refractory material with tufts of high temperature fibers anchored in it to form a pile, in the manner of a carpet. The pile is on the inside of the nozzle, so that propulsive gases of the rocket may impinge directly thereon. The exit cone may be either rigid, as is conventional, or foldable so that the rocket may be stored in a minimal space. The folded exit cone may be extended by any of various actuating systems when the rocket is launched. This pile construction may also be used for other surfaces that must be exposed to hot, high velocity gases.

Method of controlling the spin rate of tube launched rockets

Abstract: The spin rate of a rocket missile is controlled by bevelling the inside ofhe leading edge of the missile fins. Altering the bevel angle controls the rate of spin. The spin rate is also controlled by sweeping back the leading edges of the fins.

Flight tests of Viking parachute system in three Mach number regimes. 1: Vehicle description, test operations, and performance

Abstract: Flight qualifications for parachutes were tested on full-scale simulated Viking spacecraft at entry conditions for the Viking 1975 mission to Mars. The vehicle was carried to an altitude of 36.6 km for the supersonic and transonic tests by a 980.000 cu m balloon. The vehicles were released and propelled to test conditions with rocket engines. A 117,940 cu m balloon carried the test vehicle to an altitude of 27.5 km and the conditions for the subsonic tests were achieved in free fall. Aeroshell separation occurred on all test vehicles from 8 to 14 seconds after parachute deployment. This report describes: (1) the test vehicle; (2) methods used to insure that the test conditions were achieved; and (3) the balloon system design and operations. The report also presents the performance data from onboard and ground based instruments and the results from a statistical trajectory program which gives a continuous history of test-vehicle motions.

Liquid fuel multistage rocket

Abstract: This invention is a rocket which has a thrust stage located at the lowermost stage and an arbitrary number of propellant stages which are cut off in turn from a propellant stage located at the uppermost stage after exhausting fuel therein.

Means for improving rocket missile accuracy

Abstract: This invention comprises wrap-around fins for a rocket missile, the fins ng mounted on rods which are borne by a spool-like runner which encircles the nozzle body of a rocket missile. The nozzle body is exteriorly coated with teflon. The spool-like member has a ring-shaped extension on each end, each of which fits into a ring-shaped teflon bearing which, in turn, fits into a slot in a flange on the nozzle body whereby the spool-like member and the fins which are mounted on it are free to rotate relative to the nozzle (and therefore the rocket) body. The interior of the nozzle of the rocket is formed with flutes, the action of the exiting gases on the flutes causing the missile body to rotate in flight.

Mid‐latitude artificial noctilucent clouds initiated by high‐altitude rockets

Abstract: On two occasions, in February 1971 and March 1972, the exhausts produced by a rocket reentering into the atmosphere have initiated the formation of an artificial noctilucent cloud in France at an altitude of around 80 km. As the exhausts did not produce a persistent trail on the upper part of the rocket trajectory during burning, it is assumed that the cloud was of water, although an initial amount of only 260 g of water was released.

Rocket Exhaust Flow in Tube Launchers

Abstract: HE structural design of a lightweight, cylindrical-tube rocket launcher requires the calculation of the maximum pressure exerted on the tube wall by the flow exhausting from the rocket nozzle Thus, an understanding of the resultant flowfield in the launcher is needed The demand for more powerful rockets coupled with the requirement for lightweight launchers has recently emphasized the need for understanding the flowfield Test programs have been conducted to demonstrate the structural integrity of the launcher but have not used the opportunity to obtain flowfield information Contents A highly underexpanded, supersonic, short-duration jet exhausting from a conical nozzle into a short tube with an inside diameter slightly larger than the nozzle exit characterizes the flowfield of the tube-launched rockets studied The experimental investigation included eight tests in which double-base, solid-propellant rockets were statically fired into the instrumented tubes and 48 tests in which dry nitrogen was expanded through a conical nozzle into instrumented cylindrical tubes of various lengths Because the cold-gas tests formed the principal data base, the simulation of the actual rocket exhaust flows was an important consideration A full-scale facility was designed to simulate the flowfield characteristics of the rocket exhaust For both types of tests, the reservoir pressure varied rapidly with time, with the maximum value being as high as 350 atm The test times were relatively short, eg, approximately 10 msec for rocket firings and approximately 300 msec for the cold-gas tests In addition, the tube-wall pressure distribution was obtained with a rocket accelerating through a tube launcher To correlate the experimental data, a theoretical model based upon the method of characteristics was developed for the flowfield from the nozzle exit to a point just downstream of the impingement shock A sketch of the over-all flowfield based on the theory and substantiated by the acquired data is presented in Fig 1 The axially symmetric, isentropic flow region is bounded by the in viscid jet boundary and the impingement shock In actuality, this region may not be isentropic as assumed, if a strong boundary shock develops However the assumption was found to be valid for the cases considered in the present investigation Once the jet boundary intersects the tube wall an impingement shock develops as a result of the turning of the gases along the wall The flow aft of the impingement shock is highly rotational From the launcher design viewpoint, this is the most important region because maximum wall pressure exists behind the impingement shock The impingement shock projects to the tube axis as an apparent paraboloid surface If the tube is sufficiently long, the impingement shock reflects from the axis back to the wall, producing a second pressure peak Additional reflections take place depending upon the tube length Steady flow is assumed However, in reality, the flow is not steady during rocket motor ignition or chamber pressure tailoff The study concentrates on those data obtained during the "quasi-steady" flow which exists during the interval when the stagnation pressure is a maximum The flowfield during this time interval is of interest from an engineering standpoint, because the stresses in the tube wall are at a maximum

Gas Dynamic Gain of Supersonic Thrust Nozzles

Abstract: Thrust misalignment in rocket nozzles is investigated for the situation where the gas flow field is misaligned with respect to the axis of the supersonic region of the nozzle. The analysis is based upon the threedimensional method of characteristics, and hence considers the full nonlinearity of the flow field. Numerical results are presented for a wide range of nozzle geometries. These results yield considerable insight into the thrust misalignment phenomenon, and indicate means by which thrust misalignment arising from certain kinds of nozzle misalignment can be reduced. Implications to thrust vector control are also discussed.

A method for determining optimum phasing of a multiphase propulsion system for a single-stage vehicle with linearized inert weight

Abstract: A general analytical treatment is presented of a single-stage vehicle with multiple propulsion phases. A closed-form solution for the cost and for the performance and a derivation of the optimal phasing of the propulsion are included. Linearized variations in the inert weight elements are included, and the function to be minimized can be selected. The derivation of optimal phasing results in a set of nonlinear algebraic equations for optimal fuel volumes, for which a solution method is outlined. Three specific example cases are analyzed: minimum gross lift-off weight, minimum inert weight, and a minimized general function for a two-phase vehicle. The results for the two-phase vehicle are applied to the dual-fuel rocket. Comparisons with single-fuel vehicles indicate that dual-fuel vehicles can have lower inert weight either by development of a dual-fuel engine or by parallel burning of separate engines from lift-off.

ONERA and SNPE Recent Work on Solid Propellant Combustion Instability

Abstract: Results are presented on the response of a solid propellant to pressure and velocity fluctuations. Two different experimental techniques have been developed. The first makes use of devices known to be unstable (L* or T rocket motors). Comparison of cold and hot tests shows the poor credibility of the results obtained with T rocket motors; furthermore the temperature heterogeneity is scale dependent. In the second technique, research rockets with a modulated throat section are used. These methods are discussed in some detail from the experimental and theoretical standpoints. Here again cold gas comparative tests emphasize the effect of combustion on the observed gain. A linearized theory predicts the observed trends.