Showing papers on "Rocket published in 1989"

Structure of the bipolar plasma sheath generated by SPEAR I

Abstract: The Space Power Experiment Aboard Rockets I (SPEAR I) biased two 10-cm radius spheres as high as 46,000 V positive with respect to an aluminum rocket body. The experiment measured the steady state current to the spheres and the floating potential of the rocket body. Three-dimensional calculations performed using NASCAP/LEO and POLAR 2.0 show that both ion-collecting and electron-collecting sheaths were formed. The rocket body potential with respect to the ionospheric plasma adjusted to achieve a balance between the electron current collected by the spheres and the secondary electron-enhanced ion current to the rocket body. This current balance was obtained with a large ion-collecting sheath that enveloped most of the electron-collecting sheath and reduced the area for collection of ionospheric electrons. The calculated current is in agreement with the flight measurement of a steady state current of less than 1/10 A. The calculations show that the rocket body was driven thousands of volts negative with respect to the ionospheric plasma. The calculated rocket potential is within the uncertainty of that inferred from ion spectrometer data. The current flowed through the space plasma. There was almost no direct charge transport between the spheres and the rocket body.

Combined air-hydrogen turbo-rocket power plant

Abstract: A combined air-hydrogen turbo-rocket engine is disclosed having a simplified construction in which the hydrogen driven turbine is formed integrally with the rotor wheel of the axial air compressor stages. The rotor stages are located downstream of a stator vane structure and are driven by gaseous hydrogen passing across the turbine blades. The hydrogen is subsequently injected into an air duct surrounding the axial air compressor and defining an airflow path having an air inlet. The hydrogen-air mixture is ignited and the burned gases are expanded through a converging-diverging exhaust nozzle.

Aerothermochemical analysis of carbon-carbon nozzle regression in solid-propellant rocket motors

Abstract: An aerothermochemical analysis for the process of carbon-carbon composite material regression in large advanced solid-propellant rocket motors has been conducted. The analytical approach is similar in spirit to the approach of Klager, Keswani, and Kuo, with the main idea of the nozzle regression being due to the carbon chemical attack by H2O. The different steps of the work have consisted of the development and applications of several numerical codes substantiated by experimental results concerning the regression rate and the surface roughness of a carbon-carbon material. The calculated results show good agreement between measured data and the predicted regression when a flow transition is assumed, in the model, between a laminar boundary layer existing on the "smooth" virgin carbon-carbon material at the firing start and a turbulent boundary layer existing on the very rough ablative carbon-carbon surface during stabilized motor operation.

Boosted kinetic energy penetrator fuze

Abstract: A fuze for missile having a linear axis passing rearwardly through a forward penetrator, an explosive charge with the penetrator, a fuze rearward the explosive charge, a propulsion rocket rearward of the fuze, and a canister rearward of the rocket and containing a deployable parachute, the fuze containing timers for deploying the parachute from the canister a predetermined interval after release of the missile from confinement, and for igniting the rocket a predetermined interval after deployment of the parachute, a further timer for causing discharge of the explosive charge a predetermined interval after axial impact of the penetrator with a target, and a safe-arm arrangement for preventing discharge of the explosive charge prior to the impact of said penetrator.

A comparison of current-voltage relationships of collectors in the earth's ionosphere with and without electron beam emission

Abstract: Currents measurements obtained by the two sections of the Cooperative High Altitude Rocket Gun Experiment-2 sounding rocket, a mother vehicle carrying a 1-keV electron gun and a daughter vehicle tethered to the mother, are compared with the results of previous models of current collection by a charged conductor in a plasma. The current collection of the daughter vehicle, a conducting body in the ionosphere, is found to agree with the Parker-Murphy (1967) limit. The additional current collection found for the mother vehicle is attributed to beam-plasma interactions.

Combination turbojet-ramjet-rocket propulsion system

Abstract: A propulsion system for a reusable spacecraft is disclosed having turbojet,amjet and rocket modes of operation. Hydrogen or exhaust gases from a gas generator drives a gas turbine which powers an air compressor in the turbojet mode. An injection device injects hydrogen and exhaust from the gas driven turbine in the combustion chamber in the turbojet mode. In the ramjet mode, only hydrogen is injected into the combustion chamber. In the ramjet mode, hydrogen and oxygen are supplied to the rocket motor. An adjustable nozzle is provided to form a variable throat convergent-divergent nozzle in the turbojet and ramjet modes and to form a divergent nozzle in the rocket mode.

Embedded pressurization system for hybrid rocket motor

Abstract: A solid propellant component grain is supported in a combustion chamber. A liquid propellant component container is mounted forward of the combustion chamber. The liquid propellant component is supplied through conduits from the container into the combustion chamber, and ignited to form combustion gas which is discharged out the rear of the combustion chamber to generate thrust for propelling a rocket. A high pressure tank containing a non-flammable gas such as helium is at least partially embedded in the grain. A conduit leads from the tank into the container, such that the high pressure gas pressurizes the container and urges the liquid propellant component to flow from the container into the combustion chamber. The tank provides internal structural support for the grain, with the wall of the combustion chamber constituting a safety barrier in the event of structural failure of the tank after the tank is filled with high pressure gas.

Hybrid rocket combustion enhancement

Abstract: The combustion of a hybrid engine is improved by continuously injecting into a precombustion chamber a hypergolic fluid such as triethyl aluminum which exothermically reacts with and vaporizes the oxidizer such as liquid oxygen. The prevaporized oxidizer evenly combusts the solid propellant grain to develop thrust.

Numerical study of turbulent reacting flows in solid-propellant ducted rocket combustors

Abstract: A theoretical model has been developed to investigate turbulent mixing and combustion processes in the main combustion chamber of a solid-propellant ducted rocket. The formulation is based on Favre-averaged conservation equations with a two-step chemical reaction scheme and is solved by a semi-implicit finite-difference method. Turbulence closure is achieved using a well-known k-e two-equation model. Calculated flow structures show good agreement with preliminary experimental results obtained from the schlieren flow-visualization study. The influences of various parameters, including dome height and inlet flow angle, on the propulsive performance of the system are investigated in detail.

Thermal insulators for rocket motors

Abstract: A thermal insulator for a rocket motor body (3) comprises a member adapted to be fitted as a sleeve on the outer surface of the body the member (3) comprising a composite of cork (7) having on its outer surface a layer of fibre reinforced polymeric material (9) wherein a substantial proportion of the fibre reinforcement comprises fibres of low-conducting material

Full-Scale Aircraft Impact Test for Evaluation of Impact Forces, Part 1: Test Plan, Test Method, and Test Results

Abstract: One of the factors considered in the design of critical concrete structures is the estimation of the global elasto-plastic structural response caused by the accidental impact of an aircraft. To estimate the response of the structure, the impact force (the force versus time relationship) must be known. Previous analytical studies have derived the forcing function using the impact velocity of the aircraft and the calculated mass and strength distribution of the aircraft. This paper describes a test conducted on April 19, 1988, at an existing rocket sled facility at Sandia National Laboratories in Albuquerque, New Mexico, USA, in which an actual F-4 Phantom aircraft was impacted at a nominal velocity of 215 m/s into an essentially rigid block of concrete. This was accomplished by supporting the F-4 on four struts that were attached to the sled track by carriage shoes to direct the path of the aircraft. Propulsion was accomplished by two stages of rockets. The concrete target was 'floated' on a set of air bearings. Data acquisition consisted of measurements of the acceleration of the fuselage and engines of the F-4, and measurements of the displacement, velocity and acceleration of the concrete target. High-speed photography recorded the impact processmore » and also permitted the determination of the impact velocity. This paper describes the test plan, method and results, while a companion paper discusses the analyses of the results. 6 refs., 11 figs.« less

A theoretical evaluation of laser-sustained plasma thruster performance

Abstract: An extensive numerical experiment has been conducted to evaluate rocket thruster performance using a laser-sustained hydrogen plasma as the propellant. The plasma was sustained using a 30 kW CO2 laser beam operated at 10.6 microns focused inside the thruster. The steady-state Navier-Stokes equations coupled with the laser power absorption process have been solved numerically. A pressure based Navier-Stokes solver using body-fitted coordinate was used to calculate the laser-supported rocket flow which included both recirculating and transonic flow regions. The local thermodynamic equilibrium (LTE) assumption was used for the plasma thermophysical and optical properties. Geometric ray tracing was adopted to describe the laser beam. Several different throat size thrusters operated at 150 and 300 kPa chamber stagnation pressure were studied. It was found that the thruster performance (vacuum specific impulse) was highly dependent on the operating conditions, and a properly designed laser supported thruster can attain a specific impulse around 1500 secs. The heat loading on the thruster wall was also estimated and was in the range of that for a conventional chemical rocket.

Propellant formulation and process

Abstract: The present invention relates to metal filaments for use as fuel additives for rocket propellants, explosives, and other pyrotechnic devices. Preferred filaments are those such as zirconium, niobium and titanium (and alloys thereof) which have very high heat of combustion.

Rocket flight direction control system

Abstract: To quickly change rocket flight direction immediately after a rocket has been launched without increasing the wing area, the rocket flight direction control system comprises four steering wings for controlling rocket flight directions; four deflectable nozzles for jetting combustion gas backward to generate rocket thrust; a controller for generating steering control signals; and at least one actuator for actuating the deflectable nozzles in response to the steering control signals in synchronism with steering motion of the steering wings. Further, when the deflectable nozzle is divided into two, fixed and movable, nozzles, the diameter of the outlet end of the fixed divergent nozzle of the overexpansion type nozzle in matched with that of the movable nozzle for prevention of gas leakage through a gap between the two nozzles.

Bipropellant rocket engines

Abstract: A construction for the delivery of propellant constituents to the main combustion chamber of a hypergolic liquid bipropellant rocket engine having a staged combustion cycle incorporating at least one precombustor comprises: a mixer for mixing together prior to delivery to the main combustion chamber a first propellant constituent comprising oxidant together with exhaust gas from the precombustor injector means for injecting the mixture of the oxidant and exhaust gas provided by the mixer into the main combustion chamber in such a manner that the injected mixture forms a recirculation zone inside the main combustion chamber, an inlet into the main combustion chamber for a second propellant constituent comprising fuel, and a delivery channel to the said inlet, the said inlet and delivery channel being disposed laterally of the injector means relative to the main direction of flow through the injector means in such a position that fuel delivered into the main combustion chamber thereby is delivered into the said recirculation zones.

Rocket Motor Service Life Calculations Based on the First-Passage Method

Abstract: The paper discusses the uses of the first-passage method in service life predictions for solid propellant rocket motors. These motors, when stored at a site for a long period of time, are subjected to environmental temperature variations, aging, and cumulative damage. Mechanical properties are considered as statistically variable quantities. Temperature variations treated alternatively as Poisson and Markov processes are compared under cold and warm climates. As expected, service lives of motors were found to be shorter at the cold site. The Markov model gives somewhat more conservative results than the Poisson model.

Rocket observations of the precipitation of electrons by ground VLF transmitters

Abstract: Recent results obtained with electric and magnetic receivers aboard a NASA sounding rocket launched on July 31, 1987 are presented which relate multiple electron spectral peaks observed in the bounce loss cone fluxes to the resonant interaction of electrons with VLF waves from ground transmitters. The correlation of transmitter signals passing through the ionosphere with the precipitated electrons was investigated. The analysis of these in situ wave and particle data addresses the propagation of waves through the ionosphere, and, through an application of the resonant theory, enables an estimation of the cold plasma density in the interaction region.

A numerical study of chemically reacting flow in nozzles

Abstract: The space station uses small rocket motors, called thrusters, for orientation control. Because of the lack of viable design tools for small rockets, the initial thruster design was basically a very small version of a large rocket motor. Thrust measurements of the initial design were lower than predicted. To improve predictions it was decided to develop a verison of the RPLUS2D reacting flow code for thruster calculations. RPLUS2D employs an implicit finite volume, lower-upper symmetric successive overrelaxation (LU-SSOR) scheme for solving the complete two-dimensional Navier-Stokes equations and species transport equations in a coupled and very efficient manner. The combustion processes are modeled by a 9-species, 18 step finite-rate chemistry model, and the turbulence is simulated by a Baldwin-Lomax algebraic model. The code is extended to handle multiple subsonic inlet conditions where the total mass flow is governed by conditions calculated at the thruster-throat. Results are shown for a thruster design where the overall mixture ratio is hydrogen rich. A calculation of a large area ratio divergent nozzle is also presented.

Rocket Propellants from Martian Resources

Abstract: In order for extensive round trip travel to become feasible in terms of mass required in low Earth orbit, propellant manufacturing at Mars becomes essential. Martian resources lend themselves to relatively easy generation of several promising propellant combinations. Not only manned missions but also smaller unmanned missions such as sample return may benefit from this technology.

Analytical modeling of effects of wires on solid motor ballistics

Abstract: Metallic wires have been employed in numerous end-burning solid-propellant rocket motors to provide burning-rate amplifications required for certain applications. These wires provide such amplification by locally augmenting heat feedback from propellant combustion products to unburned solid material, with resultant development of cones (and consequent increase in surface area). A model of this process has been developed and tested against an extensive systematic strand-rate data base, with excellent agreement between predicted and measured effects of wire type, wire diameter, and pressure on burning rates. This model is capable of treating unsteady-state phenomena and effects of gaps between wire and propellant, possibly caused by partial wire unbonds from surrounding propellant. The model has been coupled with a chamber ballistic analysis and a geometrical analysis as regards cone shape development to permit prediction of pressure-time histories in wired motors, with gap effects included. A description of the model, discussion of its calibration against the existing wired strand data base, and discussion of effects of various gap distributions on strand burning rates and motor pressure are presented. The model offers a possible explanation for anomalous effects sometimes observed in testing of high L ID (length/diameter) wired motors at low temperature subsequent to temperature cycling.

Sky platform for rocket launching

Abstract: A launching platform for rockets or space shuttle above the Earth surface. Launching Platform is carried by one or more airships, and is intended to carry a rocket or space shuttle to an altitude of 50 km, and then launch the rocket or space shuttle with satellite into space.

The acceleration and the deceleration of the tumbling period of Rocket Intercosmos 11 during the first two years after launch

Abstract: We have analysed the development of the tumbling period of Rocket Intercosmos 11 (1974-34-B) during the first 2 years after launch. We interpret the period increase, observed from August 1974 to June 1976, as being caused by torque moments due to eddy currents induced in the hollow cylinder by the magnetic field of the Earth. The spin-decay time of 1974-34-B was 1.13 yr. This compares well to results derived by Williams and Meadows in 1978 for other Soviet rocket bodies. The tumbling acceleration of 1974-34-B, observed in June 1974, is interpeted as outgassing effect of rest propellant which remained inside the rocket after burn-off. A model of the outgassing acceleration is developed and compared to the period measurements of 1974-34-B. A reasonable good agreement between observed and predicted periods can be derived by using a nonlinear regression fit. An initial mass ratio of the rest propellant and the empty rocket cylinder is estimated.

Coaxial injector for rocket combustors

Abstract: Coaxial injector for rocket combustion chambers having a combustion chamber pressure of 5 to 25 bar and for operation with two hypergolically reacting propellants, with a central body (2) for the oxidiser, the flow channel of which has a swirl body (4) and a conically widening outlet (5) with a sharp opening edge (6), and with a sleeve for the admixture of the fuel which surrounds the central body (2) concentrically in the outlet region, projects towards the combustion chamber - forming a prereaction chamber (10) - and has lateral feed openings (8) and a central channel (9) of circular-cylindrical shape in the region of the prereaction chamber. The ratio of the length L to the diameter D of the prereaction chamber is 0.4 to 0.8. Upon entry into the prereaction chamber, the speed of each propellant component is at least 5 m/s and at most 12 m/s, and the ratio of the speed of the fuel to the speed of the oxidiser is 0.7 to 1.3.

Optimal guidance and propulsion control for transatmospheric vehicles

Abstract: Problems associated with on-board trajectory optimization and with the synthesis of guidance laws are addressed for ascent to LEO of an air-breathing, single-stage-to-orbit vehicle. A multimode propulsion system is assumed which incorporates turbojet, ramjet, scramjet, and rocket engines. An energy-state approximation is applied to a four-state dynamic model for flight of a point mass over a spherical nonrotating earth. An algorithm for generating fuel-optimal climb profiles is derived via singular perturbation theory. This algorithm results from application of the minimum principle to a low-order dynamic model that includes general functional dependence on angle of attack and a normal component of thrust. Switching conditions are derived which, under appropriate assumptions, govern optimal transition from one propulsion mode to another. The use of bank angle to modulate the magnitude of the vertical component of lift is shown to improve the index performance. Numerical results illustrate the nature of the resulting fuel-optimal climb paths.