Showing papers on "Rocket published in 1986"

Vortex driving mechanism in oscillatory rocket flows

Abstract: Several mechanisms that cause hydrodynamic instability of sheared regions of flow in a rocket chamber to drive acoustical oscillations are integrated into a comprehensive model for the vortex shedding phenomenon in this analysis. The energy method is utilized to enable later extension of the linear theory described in this paper to finite amplitude problems in which nonlinear effects are likely to be dominant. A detailed model of the feedback effect that links the acoustical and vortical fields in the vicinity of the shear layer origin is presented. This is a vital element since it provides the phasing and also the strength of the vortical waves relative to the superposed acoustic field. The most important mechanism involved in the transfer of vortical energy into the acoustic field is the interaction of the vortices with a solid surface of impingement at a suitable distance and location downstream from the shear layer origin. The energy transfer in this case is equivalent to a dipole sound source, and a more efficient mechanism than the volume quadrapole effects that have previously been evaluated. Two useful analytical tools emerge. The first is a simple rule by which designers can make the necessary changes in internal ballistics to eliminate or reduce vortex driven oscillations. The second is a growth rate model that allows a quantitative assessment of vortical interactions and is compatible with the standard acoustic stability assessment procedures in widespread use.

Horizontal-takeoff transatmospheric launch system

Abstract: This invention is directed toward providing a transatmospheric launch system that is essentially totally reusable, provides wide flexibility in choice of orbit, and may be launched quickly on short notice. The system of the invention is a two-stage horizontal takeoff and landing system. An orbiter vehicle (50) is integrated into the underside of an aircraft (2). Aircraft (2) has a cavity (4) opening aftwardly and downwardly to receive vehicle (50). Vehicle (50) and aircraft (2) are releasably connected by struts (30, 32). Aircraft (2) and vehicle (50) proceed to staging conditions under air breathing and then rocket power. Rocket engine (22) of aircraft (2) is throttled to produce a thrust differential with rocket engine (66) of vehicle (50). This differential causes vehicle (50) to automatically pivot away from aircraft (2) on struts (30, 32). After pivoting out of cavity (4), vehicle (50) is disengaged from struts (30, 32) and proceeds on its own to orbit. Aircraft (2) makes a conventional landing. Following reentry, vehicle (50) makes an unpowered horizontal landing. Separation is accomplished at a Mach number of about 3.3. In a second embodiment, the main engine of the orbiter (50') is a scramjet (101) instead of a rocket.

Mixing, chemical reaction, and flowfield development in ducted rockets

Abstract: Calculations have been made of the three-dimensional mixing, chemical reaction, and flowfield development in a typical ducted rocket configuration. The governing partial differential equations are numerically solved by an iterative finite-difference solution procedure. The physical models include the A; ~ e turbulence model, onestep reaction, and mixing controlled chemical reaction rate. Radiation is neglected. The mean flow structure, fuel dispersal patterns, and temperature field are presented in detail for a base configuration with a 0.058 m dome height, 45-deg side-arm inclination, and with gaseous ethylene injected from the dome plate at an eccentric location. In addition, the influences of the geometrical parameters, such as dome height, inclination of the side arms, and location of the fuel injector, are studied.

Rocket staging system

Abstract: An improved rocket staging system for missiles and the like wherein a carriage borne rocket engine assembly is sequentially employed within separate, generally aligned oxidizer stages which are generally coaxially disposed about the central rocket engine and its associated carriage. A central fuel tank is surrounded by several separate, cooperating, generally ring-shaped oxidizer tanks generally coaxially disposed about the rocket periphery. A plurality of oxidizer delivery lines run through each of the outer tanks and up to the top of the fuel tank, where a flexible hose brings oxidizer down to the engine carriage. As fuel is consumed, the rocket motor carriage slides upwardly inside the fuel tank in response to thrust. When the carriage is firmly seated inside the next higher oxidizer tank and all of the propellant has been removed from the lowest tank, the lowest tank is jettisoned to discard unnecessary mass. Thus when a stage is jettisoned, its oxidizer lines disconnect from those of the next higher stage and check valves in the lower endsd of the lines in the next stage prevent significant oxidizer spillage. Oxidizer intake ports such as solenoid valves mounted on the oxidizer delivery lines in each stage are kept open in the lowermost stage and closed in all other stages to allow oxidizer to be drawn only from the lowermost tank.

Experimental thrust performance of a high-area-ratio rocket nozzle

Abstract: An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

The deflagration-to-detonation transition process for high-energy propellants - A review

Abstract: Introduction T HE deflagration-to-detonation transition (DDT) in solid energetic materials has been of interest to researchers for many decades since it has a variety of areas of applicability, ranging from industrial to military. A recent example of the former is the detonation that occurred during the production of propelling charges for hunting ammunition. In the military area, the applicability extends from gun systems and projectile impact hazards to rocket motors. Hence, the DDT process has been investigated for both voidless (cast) and porous systems. We shall use the rocket motor environment to assess the DDT hazards associated with high-energy propellants. In the area of solid propellant rocket motors, an oft-asked question is "Will a cast, well-manufactured rocket propellant grain undergo a transition to detonation from the burning mode?" The answer is no, to the best of our knowledge. Although there are few journal articles directly providing this assessment, our knowledge of the DDT mechanism in gases, liquids, and solids provides a rationale for reaching this conclusion. The rationale is based on the thesis that the deflagration process must ultimately produce a shock wave to drive the system to detonation.' That is, the shock-to-detonation transition (SDT) is the final stage in any DDT process. Consequently, in the solid-propellant rocket motor situation, the confinement provided by the motor case must be sufficient to allow the pressure from deflagration to build up to a sufficiently high shock pressure to initiate the cast propellant. As will be shown below, these shock amplitudes cannot be reached for cast propellant systems confined in rocket motor cases.

Design of a rocket‐borne radiometer for stratospheric ozone measurements

Abstract: A four‐filter ultraviolet radiometer for measuring stratospheric ozone is described. The payload is launched aboard a Super‐Loki rocket to an apogee of 70 km. The instrument measures the solar ultraviolet irradiance over its filter wavelengths as it descends on a parachute. The amount of ozone in the path between the radiometer and the sun is calculated from the attenuation of solar flux using the Beer–Lambert law. Radar at the launch site measures the height of the instrument throughout its flight. The fundamental ozone value measured by the ROCOZ‐A radiometer is the vertical ozone overburden as a function of geometric altitude. Ozone measurements are obtained for altitudes from 55 to 20 km, extending well above the altitude range of balloon‐borne ozone‐measuring instruments. The optics and electronics in the radiometer have been designed within relatively severe size and weight limitations imposed by the launch vehicle. The electronics in the improved rocket ozonesonde (ROCOZ‐A) provide essentially drift‐free outputs throughout 40‐min ozone soundings at stratospheric temperatures. The modest cost of the payload precludes recovery and makes the instrument a versatile tool compared to larger ozonesondes.

Finite-element application to rocket nozzle aeroelasticity

Abstract: The trend in rocket technology is toward thinner, larger expansion ratios and more flexible exit cones. Hence, susceptibility to aerodynamically induced flutter has been and is a concern to both designer and analyst. The objective of this investigation is to develop an analytical tool for treating the aeroelastic stability of thin, truncated conical shells subjected to internal supersonic flow, with particular application to rocket nozzle structures. Stiffness, mass, and damping matrices are derived for an axisymmetric conical shell frustum finite element. The matrix equation of motion for a prescribed circumferential harmonic is developed and an eigenproblem formed from which flutter instability is deduced by tracking complex eigenvalue part variation with increasing dynamic pressure. Several check problems are considered to validate the various aspects of the analytics and computer code. Following this, the flutter analysis of a gas-deployed skirt, a typical rocket nozzle element, is presented to illustrate its application to rocket nozzle hardware. For this particular case, the analytical results indicate flutter at a motor chamber pressure well above the operating chamber pressure.

Velocity controller for ramjet missile and method therefor

Abstract: A velocity controller for a ramjet missile, having a supersonic inlet proximate the peripheral skin thereof for admitting air to a combustion zone of a ramjet engine, is comprised of a variable pitch cover disposed in pivotable engagement within the inlet and an actuator in operative engagement with the cover for adjustably positioning same over an angular range and thereby modulating airflow for the purpose of controlling flight characteristics and, principally, velocity of the missile. A sensing system is provided for detecting a dynamic flight parameter indicative of velocity of the missile and generating an output characteristic thereof for controlling the actuator and, in turn, the pitch of the cover. Methods for improving the flight performance of both solid fuel ramjet missiles and ducted rocket missiles are also disclosed herein.

Rocket-type line throwing apparatus

Abstract: A line throwing rocket arrangement is disclosed which is normally stored in a generally rectangular case. The set includes at least one rocket provided with a guide sleeve generally integral with the body of the rocket but disposed exteriorly thereof. During the launching, the sleeve slides over a straight guide rod. The storage case is provided with suitable means for securement of the guide rod to the box at different angles so that different ranges can be covered by the rocket in rescue or the like operations can be covered. The advance in the art is primarily in the simplified arrangement which facilitates the operation and reduces the manufacturing costs of the set.

Method of making a flexible bearing

Abstract: An improved laminated flexible bearing wherein alternate layers of elastomer and reinforced plastic shims of the lamination are cured simultaneously with heat and pressure and the lamination is bonded to forward and aft end rings for facilitating attachment of a thrust nozzle to a rocket case.

ELF waves and ion resonances produced by an electron beam emitting rocket in the ionosphere

Abstract: Results are reported from the ECHO-6 electron-beam-injection experiment, performed in the auroral-zone ionosphere on March 30, 1983 using a sounding rocket equipped with two electron guns and a free-flying plasma-diagnostics instrument package. The data are presented in extensive graphs and diagrams and characterized in detail. Large ELF wave variations, superposed on the strong beam-sector-directed quasi-dc component, are observed in the 100-eV beam-induced plasma when the beam is injected in a transverse spiral, but not when it is injected upward parallel to the magnetic-field line. ELF activity is found to be suppressed whenever the rocket passed through field lines with auroral activity, suggesting that the waves are produced by the interaction of the beam potentials, plasma currents, and return currents neutralizing the accelerator payload.

Low-thrust rocket trajectories

Abstract: The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report.

Experimental investigation of combustion efficiency of air-augmentedrockets

Abstract: This paper reports the results of an experimental study concerned with the mixing and combustion processes in air-augmented rockets. A combustor assembly consisting of the primary rocket and the secondary, constant-area combustion chamber were utilized to simulate a typical tactical missile. Several different configurations of the primary chamber rocket nozzles have been designed and tested in order to evaluate the mixing in the afterburning combustion chamber. On the basis of the experimental results, the secondary combustor using the subsonic primary nozzle with multiple jets has been shown to achieve the highest combustion efficiency, about 30% greater than that of the single-sonic or supersonic nozzles. Instrumentation was also provided for measurements of the pressures and velocities of the ducted flow. These measurements provide a basis for an understanding of the combustion process.

Method of undetachably fastening the injection nozzles in the injection head of rocket combustion chambers

Abstract: A method of undetachably fastening the injection nozzles in bores in the injection head of rocket combustion chambers by means of diffusion welding which comprises firmly arranging the injection nozzles before the diffusion welding with a press fit in surface-sealed bores in the injection head by means of applying an intermediate layer, necessary for the subsequent diffusion welding, to make the outside diameter of the injection nozzles have the required oversize for producing the compressive stress, and by means of heating the injection head to the fitting temperature, which allows the injection nozzles to be introduced into the bores, and then comprises heating the injection head with the injection nozzles to the temperature necessary for the diffusion welding, whereupon the still hot injection head with the injection nozzles is rapidly cooled to room temperature.

High strain capability ballistic test device for solid propellant rocket motors

Abstract: A ballistic test device for simulating propellant and/or bond and/or strain conditions in a high pressure, high strain rate rocket motor The device uses internal compliant sleeves which allow radial expansion or growth of propellants, liners and insulation or any combination of all three

Automatically woven three-directional composite structures

Abstract: Three-directional (3-D) fiber reinforced composites were demonstrated with advantages for certain missile and space structures. The applications range from carbon-carbon (c-c) to carbon-epoxy structures. 3-D carbon fiber preforms were woven using automated techniques developed by Aerospatiale of France and then impregnated and processed into c-c or carbon-epoxy structures. Demonstrated structures include c-c ITEs and exit cones for rocket nozzles and carbon-epoxy adapter rings for rocket cases. Other potential applications, including satellite truss joints and meteroid impact shields for space station components, are identified. Advantages of these structures include automated fabrication, improved mechanical properties, and greater reliability. 16 figures, 1 table.

Boost-sustain-boost rocket

Abstract: A propellant configuration for a solid propellant rocket motor, comprising a generally cylindrical housing containing a generally cylindrical main grain having a central cavity, and an igniter assembly supported in the central cavity and having means therein for bringing about combustion. At the time of actuation of the igniter assembly, hot gases are directed along flutes formed in the interior of the main grain, to cause burning thereof to take place. Advantageously, the propellant utilized in the main grain is the same as utilized in the igniter assembly, with the burn surface area of the igniter grain being in a pre-established relationship to the burn surface area of the fluted main propellant grain, such that the igniter grain will be totally consumed at the time that the flutes of the main propellant are consumed. In this way the total burn area of the main grain is reduced to a value which produces a distinct and highly desirable reduction in the total mass flow rate, this preferably occurring adjacent the end of the launch mechanism.

Rocket firing system for sequential firing of rocket motor groups

Abstract: An aerial target vehicle has an aerodynamically stabilized body carrying plural rocket motors An ignition system for the motors serves to fire the motors in selected groups at selected time intervals The ignition system uses a capacitor as its sole power source The capacitor is charged only after the motors are in place and the vehicle is ready for launch A capacitor-powered logic circuit ignites the motor groups at the selected time intervals The sequential firing of the rocket motors provides a quasi-maneuverable target with relatively long range and flight time The capacitive power arrangement is a safety factor that prevents premature firing of the rocket due to transient currents

Acceleration effects in solid propellant rocket motors

Abstract: The performance variations due to acceleration loads imposed on spinning solid propellant rocket motors are investigated. The four potentially most significant modes of acceleration-induced phenomena are identified from a study of the literature and modeled. The four modes are a mechanical mode which deals with deformations of the propellant and case: a thermodynamic mode which covers acceleration-induced combustion phenomena; a stress mode which covers the stressed propellant's effect on burn rate; and a gas dynamic mode which deals with changes in gas flow in the chamber and through the nozzle. Simplified models of each mode are developed or taken from the literature and are added to an internal ballistics evaluation computer program. The resulting analysis is the first to include all of the modes. In order to do this an original analysis of the mechanical and stress modes was necessary. However, the analysis shows that the stress mode is not important for the circular perforated grains studied. The other effects are shown to have a significant influence on solid rocket motor performance. The magnitude of the different mode effects are such that one may not be ignored over the others as has been done in the past. The results of the analysis are compared to published rocket motor data. The comparisons indicate an erosive burning effect that is a function of spin rate. A qualitative explanation of the erosive effect is presented.

Three-dimensional refraction/diffraction of electromagnetic waves through rocket exhaust plumes

Abstract: An analytical model has been developed to define the adverse effects that missile exhaust plumes have on radio communications with the missile. The computational model, RD3D, describes the three-dimensional propagation of electromagnetic waves through an ionized plume by refraction, the formation of blackout or dead zones, and dual-path contributions by diffraction from the fields outside the dead zone to the missile. The model relies on highly refined computations of the nozzle and plume chemistry and flowfields to establish the electron density distributions in the plume. These distributions are collapsed to a closed-form plume model for use in RD3D, a technique that contributes to computational time savings and increased accuracy. Results show the distortion of an electromagnetic wave as it encounters the plume and the received signal strength at the missile antenna. In order to verify the model and the individual components, comparisons are made to a number of classical refraction and diffraction cases, subscale missile diffraction tests, and full-scale missile development flight data. Results indicate refraction plays only a minor role in the received signal strength, while diffraction around an opaque circular disk may closely approximate the final results for motor propellants having high alkali content.

Acoustic oscillatory pressure control for solid propellant rocket

Abstract: The periodic shedding of vortices produced in a highly sheared gas flow, such as that encountered at the grain transition boundary in a solid propellant rocket motor or at the combustor inlet to a ramjet engine, is a significant source of acoustic instability which may result in unstarting of the ramjet engine or excessive vibration of the rocket motor. By restricting the transition boundary or combustor inlet at the sudden expansion dump plane (5), such as by locating an orifice plate (13) at the dump plane, the gas flow is separated upstream and produces a vena contracta downstream of the orifice, inhibiting the feedback of acoustic pressure to the point of flow separation and preventing the formation of organized oscillations. Incorporation of such a simple thin plate orifice at the dump plane reduces the periodic acoustic oscillations and thereby controls the detrimental oscillatory pressure fluctuations.

On numerical modelling of embedded subsonic flow

Abstract: An analytical model has been developed for computing embedded subsonic flow in rocket plumes from underexpanded axisymmetric supersonic nozzles. Numerical procedures based on the analysis have been incorporated in a simplified, non-reacting exhaust structure program and calculations for representative plume conditions performed. The technique is numerically stable and has provided satisfactory predictions of Mach–disc associated embedded subsonic flow.

Effects of geometry and materials on low cycle fatigue life of turbine blades in LOX/hydrogen rocket engines

Abstract: This paper presents the results of an advanced turbine blade test program aimed at improving turbine blade low cycle fatigue (LCF) life. A total of 21 blades were tested in a blade thermal tester. The blades were made of MAR-M-246(Hf)DS and PWA-1480SC in six different geometries. The test results show that the PWA-1480SC material improved life by a factor of 1.7 to 3.0 over the current MAR-M-246(Hf)DS. The geometry changes yielded life improvements as high as 20 times the baseline blade made of PWA-1480SC and 34 times the baseline MAR-M-246DS blade.

Blowdown pulser design criteria for solid-propellant rockets

Abstract: Pulsing solid-propellant rocket motors is often employed to evaluate the stability of the combustion pressure. The time constant of the exponentially decaying oscillations generated by the pulser provides a quantitative measure of the linear stability of the combustion pressure. Experience has shown that current design methods are inadequate for estimating the pulser charge and action time required to generate a specified pulse amplitude. This paper describes analytical and experimental studies to develop improved design methods. The linearized equations for the transient ballistics of the combustion chamber and the blowdown pulser have been solved using a Fourier transform approach. The analysis shows the exponential decay time for a blowdown pulser should be approximately twice the frequency of interest. The analysis also shows the initial pulse amplitude scales linearly with the ratio of initial pulser mass flow to motor mass flow. To generate an 8°7o peak-to-peak pulse amplitude requires that this ratio be approximately unity. Cold-flow experiments show the observed trends of pulse amplitude with pulser mass flow and blowdown time agree reasonably well with the predicted trends. However, the observed pulse amplitudes are slightly higher than the predicted amplitudes.