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Showing papers on "Rocket published in 1977"


Journal ArticleDOI
TL;DR: The results of a coordinated auroral experiment involving the Atmosphere Explorer C satellite and a sounding rocket are reported in this article, where the coordinated measurements are used to infer vertical fluxes of ionization and of electron thermal energy at high altitudes.
Abstract: Results of a coordinated auroral experiment involving the Atmosphere Explorer C satellite and a sounding rocket are reported. Auroral primary electron fluxes and neutral gas densities measured by instruments on the satellite are used in a model calculation of the thermospheric manifestation of the aurora. There is encouraging agreement between the calculated and measured electron density, electron temperature, secondary electron flux, and O I emissions at 5577 and 6300 A. A discrepancy between the calculated and the rocket-measured 3914-A emission profile is discussed in terms of experiment geometry and auroral physics. The coordinated measurements are used to infer vertical fluxes of ionization and of electron thermal energy at high altitudes

134 citations


Journal ArticleDOI
TL;DR: In this paper, a fluid mechanical model was developed to assess the performance of a rocket that is propelled by the absorption of radiant energy from a remotely stationed, repetitively pulsed laser.
Abstract: A fluid mechanical model is developed to assess the performance of a rocket that is propelled by the absorption of radiant energy from a remotely stationed, repetitively pulsed laser. The model describes the flow within a conical nozzle that is subjected to point energy depositions at the apex of the cone. A similarity solution is obtained and the specific impulse and energy efficiencies that may be achieved with such a device are determined. Fluid mechanical constraints limit the range of pulse repetition rates that may be utilized. Preliminary design considerations indicate that a specific impulse of 800 sec or greater may be achieved with both a laboratory and a full-scale device. A two pound laboratory rocket can be accelerated at 10 #'s with a 15 joule laser pulsed 25,000 times per sec. A one ton rocket will require a megajoule laser operating at 350 pulses per sec to achieve an equivalent acceleration. A,a D* dm

64 citations


Journal ArticleDOI
TL;DR: In this article, a comparison of neutral wind measurements made by rockets with ionospheric drift observations made by the partial reflection drift method was made. But the drift experiment was located at the Woomera rocket range (31°S, 136°E), the site being chosen to provide as good a spatial overlap as possible with wind determinations by various rocket techniques.

46 citations


Journal ArticleDOI
TL;DR: The relatively low frequency of particle sightings from Skylab, coupled with improvements in Orbiter venting techniques, indicates that sightings of particles 2 microm and larger in radius will not seriously hamper tele cope performance provided that liquid vents and rocket firings are properly restricted.
Abstract: A sensitive IR telescope on the Space Shuttle Orbiter will be limited in its performance by fluctuations in the IR radiation from the natural environment and the contaminant atmosphere. Models of the Orbiter's contaminant atmosphere were used to predict its spectral radiance from 3 to 300 microns. At 350 km, statistical fluctuations in the radiation from a water vapor column, and a noise equivalent power were measured. This noise is somewhat smaller than the expected contribution from zodiacal light from 5 to 30 microns. The column density of all IR emitting molecules can be kept low only if restrictions on rocket firings and liquid vents are maintained. The relatively low frequency of particle sightings from Skylab, coupled with improvements in Orbiter venting techniques, indicate that sightings of particles 2 microns and larger in radius will not seriously hamper telescope performance provided that liquid vents and rocket firings are properly restricted.

38 citations


Journal ArticleDOI
TL;DR: In this paper, the development of a hybrid rocket motor based on highly aluminized fuels and FLOX oxidizers is described, and the interrelationship between aluminum loading and oxidizer composition and motor configuration on combustion efficiency is given.
Abstract: The development of a hybrid rocket motor based on highly aluminized fuels and FLOX oxidizers is described. By using fluorinated oxidizers, volatile aluminum fluorides are formed as the main combustion product, which results in a higher combustion efficiency of the aluminized fuels than with oxygen-based oxidizers. Critical motor components such as oxidizer injector, mixing diaphragm, reaction chamber, and nozzle were developed to withstand the high temperature and the extremely corrosive combustion products. Correlations between the fuel regression rate and oxidizer mass flux are obtained. The interrelationship between aluminum loading and oxidizer composition and motor configuration on combustion efficiency is given. exp th a Ae/At Af f*ex CTPB F FLOX-90

36 citations


Proceedings ArticleDOI
01 Jan 1977
TL;DR: In this article, a derivation of an explicit solution to the two point boundary-value problem of exoatmospheric guidance and trajectory optimization is presented, and a form of this algorithm has been chosen for onboard guidance, as well as real time and preflight ground targeting and trajectory shaping for the NASA Space Shuttle Program.
Abstract: A derivation of an explicit solution to the two point boundary-value problem of exoatmospheric guidance and trajectory optimization is presented. Fixed initial conditions and continuous burn, multistage thrusting are assumed. Any number of end conditions from one to six (throttling is required in the case of six) can be satisfied in an explicit and practically optimal manner. The explicit equations converge for off nominal conditions such as engine failure, abort, target switch, etc. The self starting, predictor/corrector solution involves no Newton-Rhapson iterations, numerical integration, or first guess values, and converges rapidly if physically possible. A form of this algorithm has been chosen for onboard guidance, as well as real time and preflight ground targeting and trajectory shaping for the NASA Space Shuttle Program.

34 citations


Journal ArticleDOI
TL;DR: In this article, the primary electron spectrum and the energy flux on the field lines containing auroral light in the E region were collected by electron detectors aboard a sounding rocket and compared to calculations based on spectroscopic measurements of the auroral lines 4278, 5577, and 6300 A used in predicting the energy influx and the characteristic energy of an assumed Maxwellian primary electrons spectrum for two auroral displays.
Abstract: Data are collected by electron detectors aboard a sounding rocket measuring the primary electron spectrum and the energy flux on the field lines containing auroral light in the E region. These data are compared to calculations based on spectroscopic measurements of the auroral lines 4278, 5577, and 6300 A used in predicting the energy influx and the characteristic energy of an assumed Maxwellian primary electron spectrum for two auroral displays. Data were also collected by photometers sampling the auroral light from the E region magnetically conjugate to the rocket. These data are compared to those of current ionospheric models.

31 citations


Journal ArticleDOI
01 Jan 1977
TL;DR: In this paper, a numerical solution for the spray equation has been developed for thin sprays injected into a type of stratified charge internal combustion chamber, which can treat general three dimensional geometries, using a statistical approach.
Abstract: A numerical solution technique has been developed for the spray equation and has been applied to thin sprays injected into a type of stratified charge internal combustion chamber. The difference equation method can treat general three dimensional geometries, using a statistical approach. The effects of independent variations of a large number of system parameters was studied, including initial spray dispersion, amount and type of gas swirl, gas density, injection timing, chamber geometry, initial droplet size distribution, injection velocity, drag coefficient of injection. It was found that in the particular geometry chosen the gas swirl, the droplet size distribution produced by the injector, and the chamber gas density into which the spray is injected are the most important factors influencing the spray motion and vaporization. Although the calculations reported here applied to internal combustion engine conditions, the general method is applicable to other spray-injected combustors, including stationary combustors and rocket motors. Results from calculations indicate that the technique, used in coordination with selected laboratory measurements, could significantly enhance spray-injected combustor design efforts.

29 citations


Journal ArticleDOI
TL;DR: In this paper, a physically realistic model of the effect of crossflow on composite propellant combustion, based on the bending of columnar diffusion flames by the crossflow, is presented.
Abstract: : Development of solid rocket motor designs which result in high velocity flows of product gases across burning propellant surfaces(notably, nozzleless rocket motors) is leading to increased occurrence of erosive burning. In this paper, a physically realistic picture of the effect of such crossflow on composite propellant combustion, based on the bending of columnar diffusion flames by the crossflow, is presented. This bending results in shifting of the diffusion flame heat release zone toward the surface, with consequent increased heat feedback flux from this flame to the surface and thus increased burning rate. A relatively simple analytical model based on this picture is developed for prediction of propellant burnings rate as a function of pressure and crossflow velocity, given only zero-crossflow burning rate versus pressure data. Model predictions and experimental results are compared, with reasonably good agreement being found. (Author)

29 citations


Patent
12 Dec 1977
TL;DR: A rear door for a rocket launch tube provided to prevent rocket exhaust gas flow into an empty launch tube from an associated multiple-rocket plenum chamber is described in this paper, where the door is maintained in a stored position while a missile is in the launch tube and is activated when the missile leaves the launcher.
Abstract: A rear door for a rocket launch tube provided to prevent rocket exhaust gas flow into an empty launch tube from an associated multiple-rocket plenum chamber. The door is maintained in a stored position while a missile is in the launch tube and is activated when the missile leaves the launcher. The door may be latched open and released by a sensor device at a selected position of the missile as it is leaving the launch tube. Preferably, gases from the launching missile power closure of the door, once the door is released from its open latched position. Once the door closes, a second latch locks it in place to seal off the launch tube from the plenum chamber.

25 citations


Patent
03 Feb 1977
TL;DR: In this article, an identification impedance is derived from the firing wire or lead which is also used to fire the rocket, which is used to identify the type of rocket or other ammunition loaded and ready to fire.
Abstract: The system for electrically identifying and firing ammunition is especially advantageous for firing military rockets from a multiple tube rocket launcher. The system provides an electrical signal for initially identifying the type of rocket or other ammunition loaded and ready to fire in each launching tube. The identifying signal is derived from the firing wire or lead which is also used to fire the rocket. In each rocket an igniter and a diode rectifier are connected in series in the firing circuit which usually extends between the firing lead and a common ground. The diode rectifier is polarized to conduct the firing current while being nonconductive as to the signal current which is oppositely polarized. Each rocket includes an identification impedance having a nature to identify the type of rocket. The different types of rockets have measurably different identification impedances.

Journal ArticleDOI
TL;DR: A review of the literature on platonization of double-base propellant with different inorganic and organic compounds of lead and other metals can be found in this paper, where various theories to explain the mechanism of the plateau effect are presented.
Abstract: HE inclusion of small quantities of various lead compounds in the double-base propellant system results in increased burning rate at low pressure, followed by a plateau burning region, where the burning rate remains almost independent of variation in pressure, and the postplateau region, in which the burning rate/pressure relationship is similar to that of an unleaded propellant (Fig. 1). This phenomenon of independence of burning rate to pressure, generally known as "platonizatio n," implies a low value of pressure index n in the plateau region. A low value of n permits the design of lighter rocket motors because of lower safety factors. The magnitude of n is one of the important factors in determining the suitability of a propellant for rocket propulsion applications. The importance of low-pressure ballistics was realized as early as in 1939 with the widespread use of rockets. Earlier studies centered around low-energy nitrocellulose propellants with a view toward minimizing nozzle erosion, burning rate, and pressure index. However, the real impetus to the platonization of double-base propellants came from the observation that the use of lead stearate as a lubricant in the extrusion of large rocket charges actually had modified their burning characteristics. This paper reviews the literature on platonization of double-base propellant with different inorganic and organic compounds of lead and other metals, and presents various theories to explain the mechanism of the plateau effect.

Journal ArticleDOI
TL;DR: Total solar irradiance was observed simultaneously outside the earth's atmosphere by three types of absolute cavity radiometers and duplicates of four of the Nimbus 6 Earth Radiation Budget (ERB) solar channels in a June 1976 Sounding Rocket Experiment.
Abstract: Total solar irradiance was observed simultaneously outside the earth's atmosphere by three types of absolute cavity radiometers and duplicates of four of the Nimbus 6 Earth Radiation Budget (ERB) solar channels in a June 1976 sounding rocket experiment. The preliminary average solar constant result from the cavity radiometers is 1367 Wm (-2) with an uncertainty of less than + or - 0.5% in S.I. units. The duplicate ERB channel 3 on the rocket gave a value of 1389 Wm (-2) which agreed exactly with the Nimbus 6 ERB channel 3 measurement made simultaneously with the rocket flight.

Patent
12 Dec 1977
TL;DR: In this article, a system utilizing a rocket plenum design which is of a form to reduce and control combustion therein is described, where the plenum is provided with two oppositely and upwardly extending exhaust ducts.
Abstract: A system utilizing a rocket plenum design which is of a form to reduce and control combustion therein. The plenum is provided with two oppositely and upwardly extending exhaust ducts. Provision is made to eliminate blind pockets and stagnation passages in order to prevent possible explosions in the plenum during rocket firing.

Patent
Lee F. Carey1
11 Nov 1977
TL;DR: A skirt-like device attached to the exit end of a rocket engine nozzle is fabricated of thin sheet heat resistant metal; which when extended in operative condition is in frustoconical form, but prior to engine firing is pleat-folded inwardly into a stowed position relative to the exiting end of the rocket engine as discussed by the authors.
Abstract: A skirt-like device attached to the exit end of a rocket engine nozzle is fabricated of thin sheet heat resistant metal; which when extended in operative condition is in frustoconical form, but which prior to engine firing is pleat-folded inwardly into a stowed position relative to the exit end of the rocket engine nozzle. Thus, prior to engine firing the device adds little or nothing to the engine/nozzle occupancy space; but incident to engine firing the device deploys to provide an effective engine nozzle/skirt combination of increased length and exit diameter/area; such combination having an inner surface of constant slope.

01 Oct 1977
TL;DR: In this paper, the application of dual-mode propulsion concepts to fully reusable single-stage-to-orbit (SSTO) vehicles is discussed, and the potential for significant cost and performance benefits when applied to SSTO vehicles.
Abstract: The application of dual-mode propulsion concepts to fully reusable single-stage-to-orbit (SSTO) vehicles is discussed. Dual-mode propulsion uses main rocket engines that consume hydrocarbon fuels as well as liquid hydrogen fuel. Liquid oxygen is used as the oxidizer. These engine concepts were integrated into transportation vehicle designs capable of vertical takeoff, delivering a payload to earth orbit, and return to earth with a horizontal landing. Benefits of these vehicles were assessed and compared with vehicles using single-mode propulsion (liquid hydrogen and oxygen engines). Technology requirements for such advanced transportation systems were identified. Figures of merit, including life-cycle cost savings and research costs, were derived for dual-mode technology programs, and were used for assessments of potential benefits of proposed technology activities. Dual-mode propulsion concepts display potential for significant cost and performance benefits when applied to SSTO vehicles.

01 May 1977
TL;DR: In this article, it was demonstrated that cryogenic propellants can be stored unvented in space long enough to accomplish a Saturn orbiter mission after 1,200-day coast.
Abstract: It was demonstrated that cryogenic propellants can be stored unvented in space long enough to accomplish a Saturn orbiter mission after 1,200-day coast. The thermal design of a hydrogen-fluorine rocket stage was carried out, and the hydrogen tank, its support structure, and thermal protection system were tested in a vacuum chamber. Heat transfer rates of approximately 23 W were measured in tests to simulate the near-Earth portion of the mission. Tests to simulate the majority of the time the vehicle would be in deep space and sun-oriented resulted in a heat transfer rate of 0.11 W.

Patent
20 May 1977
TL;DR: In this article, a toy including a number of press-fitting parts configurable as a rocket plane is described, including optical elements which may function as a kaleidoscope, or, in a second configuration, as a periscope.
Abstract: A toy including a number of press-fitting parts configurable as a rocket plane is provided. A number of the parts include optical elements which may function as a kaleidoscope, or, in a second configuration, as a periscope. Some of the parts may be assembled to form a blow-gun for discharging a toy projectile at a toy target.

Patent
Lee F. Carey1
24 Jun 1977
TL;DR: A skirt-like device attached to the exit end of a rocket engine nozzle is fabricated of thin sheet heat resistant metal; which when extended in operative condition is in frustoconical form, but prior to engine firing is pleat-folded inwardly into a stowed position relative to the exiting end of the rocket engine as discussed by the authors.
Abstract: A skirt-like device attached to the exit end of a rocket engine nozzle is fabricated of thin sheet heat resistant metal; which when extended in operative condition is in frustoconical form, but which prior to engine firing is pleat-folded inwardly into a stowed position relative to the exit end of the rocket engine nozzle. Thus, prior to engine firing the device adds little or nothing to the engine/nozzle occupancy space; but incident to engine firing the device deploys to provide an effective engine nozzle/skirt combination of increased length and exit diameter/area; said combination having an inner surface of constant slope.

Journal ArticleDOI
TL;DR: In this article, an experimental investigation of a simple resonance ignition tube has been conducted, and strong resonance was found when the leading edge of the resonance tube was located within the third compression cell of the underexpanded jet.
Abstract: An experimental investigation of a simple resonance ignition tube has been conducted. The leading edge of the resonance tube was blunt in order to simulate the geometry of proposed rocket ignition systems. The jetstagnation pressure, the nozzle/resonance-tube separation distance, and the resonance-tube length were varied systematically. Strong resonance was found to occur when the leading edge of the resonance tube was located within the third compression cell of the underexpanded jet. High-speed motion pictures showed a significant upstream propagation of disturbances from the resonance tube to the nozzle. These results will be useful in determining the feasibility of using resonance-tube ignition systems for both liquidand solid-propellant rocket engines.

Patent
Fred R. Youngren1
16 Feb 1977
TL;DR: In this paper, the authors describe an indicator attached to the casings of rocket motors in the field to show the readiness of a rocket motors using solid fuel using a viscous material, such as a wax, disposed in a chamber and a spring-loaded member which moves in accordance with changes in ambient temperature over a period of time.
Abstract: Forms of an indicator attached to the casings of rocket motors in the field to show the readiness of rocket motors using solid fuel are described. Each form described includes a viscous material, as for example a wax, disposed in a chamber and a spring-loaded member which moves in accordance with changes in ambient temperature over a period of time. The amount of movement is analogous to change in the characteristics of the solid fuel.

01 Oct 1977
TL;DR: In this article, a preliminary study of a space transport concept consisting of a winged orbiter containing ascent propellants and two small turbojet-powered winged boosters, used for acceleration to supersonic speeds is presented.
Abstract: Results of a preliminary study of a space transport concept are presented. The concept consists of a winged orbiter containing ascent propellants and two small turbojet-powered winged boosters, used for acceleration to supersonic speeds. The concept offers full reuse and horizontal takeoff from numerous existing airports. With current structure and rocket technology, this transport concept has lower gross weight for a selected payload than single-stage-to-orbit concepts, which require structural advancements. Discussion includes alternatives to the baseline space transport concept which improve performance; such as advanced structures, dual-fuel rockets, and lightweight scramjets for the orbiter. A concept of using a stretched shuttle orbiter instead of an all-new configuration is discussed for two payload-class vehicles.




Patent
04 Nov 1977
TL;DR: In this paper, a combustible rocket is constructed with a laminate which comprises high tensile strength metal foils, easily combustible metal foam layers, and easily combustable adhesive layers.
Abstract: A combustible rocket is constructed with a laminate which comprises high tensile strength metal foils, easily combustible metal foils, and easily combustible adhesive layers. The rocket has sufficient mechanical strength and its shell can be gasified or disintegrated into finely divided pieces after the propellant of the rocket has burnt out at high altitudes. This rocket can be launched safely even in the vicinity of population centers without risk of injury to living things.

Journal ArticleDOI
TL;DR: A Nike-Tomahawk rocket was launched into a system of auroral arcs northward of Poker Flat Research Range, Fairbanks, Alaska, and the pitch-angle distribution of electrons was measured at 2.5, 5, and 10 keV as discussed by the authors.
Abstract: A Nike-Tomahawk rocket was launched into a system of auroral arcs northward of Poker Flat Research Range, Fairbanks, Alaska. The pitch-angle distribution of electrons was measured at 2.5, 5, and 10 keV and also at 10 keV on a separating forward section of the payload. The auroral activity appeared to be the extension of substorm activity centered to the east. The rocket crossed a westward-propagating fold in the brightest band. The electron spectrum was relatively hard through most of the flight, showing a peak in the range from 2.5 to 10 keV in the weaker aurora and below 5 keV in the brightest arc. The detailed structure of the pitch-angle distribution suggested that, at times, a very selective process was accelerating some electrons in the magnetic field direction, so that a narrow field-aligned component appeared superimposed on a more isotropic distribution. It is concluded that this process could not be a near-ionosphere field-aligned potential drop, although the more isotropic component may have been produced by a parallel electric field extending several thousand kilometers along the field line above the ionosphere.

Patent
23 Mar 1977
TL;DR: In this article, an electric detonator can be used to destroy all the remaining parts of a rocket and casing, which can be remotely operated by a remote operator, e.g. a user.
Abstract: To spray crops from the air or to treat clouds chemically, or to distribute anti radar media, etc., a rocket (12) can be used. This contains a propulsive charge (124) of E.G. propergol powder, which is greater than that required to reach the largest area. This is partially surrounded by the medium (130) to be dispersed, leaving an unburnt portion (128) inside the latter charge (130). Ahead of this is a detonating charge (132) which can be fired by e.g. an electric detonator (134). In this way, the remainder (132) of the explosive charge can be ignited so as to destroy all the remaining parts of teh rocket and casing. The electric detonator can be remote operated.

Journal ArticleDOI
TL;DR: In this article, a simple methodology, amenable to hand calculation, is presented for this purpose, the user need supply only the propellant aluminum concentration, missile velocity and altitude, and the rocket motor thrust level.
Abstract: Ambient light scattered from metal oxide particles in composite-propellant rocket exhausts makes the missile trail visible at long ranges. In the selection of propellant aluminum concentration, it is important to be able to predict the distance from which the exhaust no longer is visible. In this paper, a simple methodology, amenable to hand calculation, is presented for this purpose. The user need supply only the propellant aluminum concentration, missile velocity and altitude, and the rocket motor thrust level. The methodology predicts smoke trail dimensions, particle concentration as a function of location in the trail, sunlight and skylight scattering, background obscuration, and the range at which the trail may be detected visually for any atmospheric attenuation level. The entire calculation can be carried out with a scientific pocket calculator.

01 Nov 1977
TL;DR: In this paper, composite propulsion was analyzed for single-stage-to-orbit vehicles designed for horizontal take-off, and trajectories, geometric, and mass analyses were performed to establish the orbital payload capability of six engines.
Abstract: Composite propulsion was analyzed for single-stage-to-orbit vehicles designed for horizontal take-off. Trajectory, geometric, and mass analyses were performed to establish the orbital payload capability of six engines. The results indicated that none of the engines performed adequately to deliver payloads to orbit as analyzed. The single-stage turbine and oxidizer-rich gas generator resulted in a low engine specific impulse, and the performance increment of the ejector subsystem was less than that of a separate rocket system with a high combustion pressure. There was a benefit from incorporating a fan into the engine, and removal of the fan from the airstream during the ramjet mode increased the orbital payload capability.