Showing papers on "Rocket published in 1999"

Flow Separation and Side-Loads in Rocket Nozzles

Abstract: Flow separation in nozzles of rocket engines is undesired because it can lead to dangerous lateral forces, which might damage the nozzle The origin of side-loads is not fully clear, although different possible origins were identified in the past Meanwhile, it seems to be clear that in thrust-optimized or parabolic nozzles, a major side-load is due to the transition of separation pattern from free shock separation to restricted shock separation and vice versa After a literature review, the reasons for the transition between the separation patterns are discussed, and the cap shock pattern, which is identified to be the cause of this transition, is closely analyzed It turns out that this pattern can be interpreted as an inverse Mach reflection of the internal shock at the centerline The separation and side-load behavior of thrust-optimized and parabolic nozzles is described in detail In order to be able to predict the pressure ratio pc/pa at which the transition of separation patterns occurs, a model is developed which uses TDK-data as an input With the oblique shock relations and a momentum balance, both the ratio of chamber to ambient pressure and the value of the lateral force can be predicted with fair accuracy

Friction stir welding takes off at Boeing

Abstract: The twin benefits of cost savings and better weld properties have led Boeing to switch to friction stir welding for use on its Delta rocket programm.

Payload carry and launch system

Abstract: The reusable space launch system (1) embodiment has a first stage vehicle or aerospacecraft (50), a second stage vehicle or reusable spacecraft (51) and a third stage vehicle or reusable orbit transfercraft (52). All the stages have the basic aerodynamic vehicle elements of a fuselage, wings, and tail, with the incorporation of control surfaces to supply lift, stability and control. The aerospacecraft (50) is configured to use ejector ramjet engines (18) for powered flight and includes equipment to capture air to supplement oxidizer for the ejector ramjet engine (18) during take-off and extreme high altitude. In order to optimize aerospacecraft (50) performance in a pull up movement to exit the sensible atmosphere, the aerospacecraft (50) may include auxiliary ascent rocket engines (116). The aerospacecraft (50) has a payload bay that is accessed by nose load re-closable payload fairings (74).

Flow Separation and Side-Load Behavior of the Vulcain Engine

Abstract: The Vulcain engine flow separation and side-load behavior observed and measured during thrust chamber tests is discudded in detail in this paper. It is shown by the test results and by comparison with numerical flow data that the parabolic Vulcain nozzle features a transition in separation behavior from free shock separation to restricted shock separation and vice versa during both engine start-up and shut-down. These highly transient phenomena are a major cause of side-loads. In addition, the side-load activities are measured during nozzle operation with pre free shock separation or pure restricted shock separation. By using results from numerical simulations, it is shown that a specific plume pattern, the cap-shock pattern, is responsible for the observed flow transition. Finally, a comparison of the flow behavior in the Vulcain nozzle during start-up and shut-down is compared with other published data for thrust-optimized or parabolic rocket nozzles with an internal shock emanating from the throat.

Hybrid rocket system and integrated motor for use therein

Abstract: A hybrid rocket system and motor includes an oxidant tank communicating with a combustion chamber containing a solid fuel grain. The passage from the oxidant tank to the combustion chamber is obstructed by a fill tube which fills the oxidant tank with oxidant and pressurizes it. Once the tank is full, the fill tube is displaced from the tank-chamber passage to deliver oxidant to the grain and initiate the combustion process. The motor may be an integral part of the rocket, or it may be a modular, expendable or exchangeable cartridge-type device inserted into a reusable body.

Dynamic Modeling of Hybrid Rocket Combustion

Abstract: The combustion dynamics in hybrid rockets is studied to provide a basic input into transient motor processes. The model treats the time-dependent heat e ow into the ablating fuel surface. A variable surface temperature is considered with an effective activation energy to describe the surface-temperature variation during the transient. Two time scales are observed for throttling: a short lag near the surface related to the activation energy, and the larger well-known thermal lag as a result of conductivity. The model is also applied to an oscillating surface heate ux input. Weobserved an amplie cation oftheregression-rateoscillations for low frequencies.Although thiseffect is not the cause of instability, it can aggravate existing oscillations at these low frequencies. We next formulated a quasisteady combustion model, which is then coupled with thethermal lag system with boundary-layer delaysthat account for the adjustment of the boundary layer to the changes in the freestream conditions and blowing from the surface. A linearized treatment of this coupled system evidences some low-frequency instabilities. The scaling of the oscillation frequencies and the erratic character of the experimentally observed instabilities are successfully explained.

The Physical Origins of Side Loads in Rocket Nozzles

Abstract: The side loads that typically occur during the start-up and shut-down transients of rocket nozzles are not yet a fully understood phenomenon. Despite the great practical interest to avoid structural problems to the engine, only few clean data are available, because of the difficulties and the high costs of both experimental and numerical tests. Trying to fill this lack of knowledge, previous numerical studies of transient solutions by the authors are re-examined and compared with numerical steady-state solutions at different ambient conditions. Moreover, the numerical and experimental data published in the meantime are discussed and compared with the present numerical simulations. The analysis allows to establish the main features of these flowfields, and to indicate important differences between steady and transient flow configurations.

Toy rocket launcher

Abstract: A rocket launcher ( 10 ) is provided having a manual air pump ( 11 ) coupled to a base unit ( 12 ) through an elongated pressure tube ( 13 ) having a pressure release valve or trigger ( 20 ). The base unit has a pressure chamber ( 24 ), a launch tube ( 25 ), and an orientation sensitive safety mechanism ( 32 ) coupled to the end of the pressure tube. The safety mechanism is coupled to an internal pressure tube ( 33 ) which extends to a pressure sensitive release valve ( 34 ) that controls the release of pressurized air from the launch tube. The orientation sensitive safety mechanism prevents the launching of projectiles should the launch tube be offset from a vertical orientation and depressurizes the launcher should an operator attempt to fire the launcher while in an offset orientation.

Reconstruction Technique for Reducing Hybrid-Rocket Combustion Test Data

Abstract: A new integral technique has been developed for the purpose of analyzing combustion test data in hybrid rocket firings. The method, which utilizes a complete reconstruction of the regression history, is shown to have several advantages when compared with the current endpoint technique, in which average regression rate and port mass flux are determined based on fuel grain initial and final dimensions. One significant advantage is that the integral method permits one to ascertain the sensitivity of regression rate to changes in mass flux in a single test. Action (burning) times greater than 5 s have been found to be preferred over shorter values more typically used to date, because this minimizes errors in action time determination and permits the fuel to reach a steady-state thermal profile.

Analysis of plasma behavior in a magnetic thrust chamber of a laser fusion rocket

Abstract: A magnetic thrust chamber concept in a laser fusion rocket is suitable for controlling the plasma flow, and it has an advantage in that thermalization with wall structures in a thrust chamber can b...

Some Aspects of Time‐Temperature Superposition Principle Applied for Predicting Mechanical Properties of Solid Rocket Propellants

Abstract: Optimal test conditions for determining the mechanical properties of rocket propellants (temperatures and strain rates ranges) for delivering master curves were investigated. From master curves it is possible to predict the modulus, maximum stress and maximum strain in vide intervals of temperatures and strain rates, and especially the existing conditions during the ignition of rocket motor. Using the control experiments, at high strain rates, the good agreement between the results obtained from master curves was shown. The obtained results for composite rocket propellants (with carboxy-terminated polybutadiene, CTPB, as a binder), point out the drastic decreasing of maximum strain at high strain rates and low temperatures.

Axisymmetrical annular plug propulsion system for integrated rocket/ramjet or rocket/scramjet

Abstract: The invention is an airframe which includes a vehicle ( 12 ) having a solid propellant rocket engine ( 14 ) and a ramjet or scramjet engine ( 16 ); a thrust plug ( 18 ) extending from an end ( 20 ) of the vehicle which directs combustion gases ( 23 and 64 ) produced by the solid propellant rocket engine or ramjet/scramjet engine to produce forward thrust; a longitudinal passage ( 38 ) extending from the end of the vehicle to an opening ( 30 ) forward of the end which receives external air directed by forward movement of the vehicle and in which solid propellant ( 32 ) of the solid propellant rocket engine is located, and wherein during rocket operation solid propellant is combusted to produce the combustion gases in the longitudinal passage which are conveyed by the longitudinal passage into contact with the thrust plug and during ramjet/scramjet operation the longitudinal passage is open to flow of external air after operation of the solid propellant rocket engine is completed and which supports mixing and combustion of the air/fuel by the ramjet/scramjet engine to produce combustion gases which are conveyed by the longitudinal passage into contact with the thrust plug.

High Regression Rate Hybrid Rocket Propellants and Method of Selecting

Abstract: This invention comprises a new process for developing high regression rate propellants for application to hybrid rockets and solid fuel ramjets. The process involves the use of a criterion to identify propellants which form an unstable liquid layer on the melting surface of the propellant. Entrainment of droplets from the unstable liquid-gas interface can substantially increase propellant mass transfer leading to much higher surface regression rates over those that can be achieved with conventional hybrid propellants. The main reason is that entrainment is not limited by heat transfer to the propellant from the combustion zone. The process has been used to identify a new class of non-cryogenic hybrid fuels whose regression rate characteristics can be tailored for a given mission. The fuel can be used as the basis for a simpler hybrid rocket design with reduced cost, reduced complexity and increased performance.

Aluminum combustion modeling in solid propellant environments

Abstract: An improved numerical model of aluminum particle combustion has been developed based on previous modeling work done at Brigham Young University. Thermochemical calculations have been performed to determine the range of temperatures, pressures, and species distributions resulting from the combustion of various formulations of rocket propellants. The previouslydeveloped model has been modified to take into account such species distributions at high temperatures and pressures. Kinetic rate equations are introduced for the homogeneous reaction of COZ and Hz0 with ahuninum The model has produced results which are consistent with experimental data from many authors. In addition, the model has been used to predict alumirmm combustion rates for the range of conditions predicted by thermochemical calculations to be present inside the rocket motor, for which there is little or no experimental data. The effects of less reactive species such as HCl and HZ have been included for pressures up to 70 atm and temperatures up to 4000 K. Introduction Aluminum powder is presently used, and has been used for many years, as an ingredient to increase the specific impulse of solid rocket propellants because of the large amount of heat generated during the aluminum oxidation reaction. This occurs when aluminum reacts with the available Hz0 and CO, (the major oxidizing species) produced from a burning solid propellant In solid propellants, aluminum is used in quantities of lo-20% by mass, and the particles are typically 20-30 microns in diameter. During heat-up, these particles may melt and coalesce into larger agglomerates, ranging from 100-200 microns in size. It is very useful to be able to predict the time required for the aluminum particles to burn once they are ignited. Empirical correlations may be used, but data collected from rocket motors themselves are virtuahy non-existent because of the harshness of the motor environment Most empirical correlations have therefore been derived from carefully controlled lab experiments. However, these experiments differ greatly from the conditions in a rocket motor, and while such experiments are informative, it can be difficult to extrapolate their results to the conditions of a rocket motor. Computer modeling of ahnninq combustion thus becomes an attractive alternative. Background When aluminum ignites, the heat.of reaction is so great that the aluminum boils and thus remains at about 2800K (at 1 atm). The outward flux of gaseous ahuninum causes a flame zone to form at about 2-4 times’ the diameter of the particle, in a manner similar to that of a burning hydrocarbon droplet. In this thune zone a homogeneous reaction takes place between the ahnninq and, available oxidizer(s). Figure 1 shows a representation of the vaporphase aluminum combustion process. In the flame zone, the oxidized products consist of gas-phase AlxO, species such as AlO. Figure 1. Representation of the vapor-phase aluminum combustion process.’ There is no gas phase Al203 since Al203 dissociates at its boiling point However, there is usually not enough heat for all the oxide species to remain in the gas phase, so some A&O, species condense and associate to form

Rocket trajectory estimating method, rocket future-position predicting method, rocket identifying method, and rocket situation detecting method

Abstract: There is provided a rocket trajectory estimating method comprising the steps of: measuring a GLOS angle of a flying rocket a tracking system; passing the GLOS angle data through a batch filter to reduce noises; estimating a rocket trajectory on the basis of the GLOS angle data, the noises of which have been reduced; passing the resulting rocket trajectory data through a Kalman filter to reduce biases; and estimating the rocket trajectory again on the basis of the corrected GLOS angle data and the positional information of the tracking system. Thus, there is provide a rocket trajectory estimating method capable of reducing observation errors (noises and biases) of a tracking system of a passive ranging system, which does not have need of any laser range finders, to enhance the accuracy of rocket trajectory estimation.

Analysis of a New Rocket-Based Combined-Cycle Engine Concept at Low Speed

Abstract: An analysis of the Independent Ramjet Stream (IRS) cycle is presented. The IRS cycle is a variation of the conventional ejector-Ramjet, and is used at low speed in a rocket-based combined-cycle (RBCC) propulsion system. In this new cycle, complete mixing between the rocket and ramjet streams is not required, and a single rocket chamber can be used without a long mixing duct. Furthermore, this concept allows flexibility in controlling the thermal choke process. The resulting propulsion system is intended to be simpler, more robust, and lighter than an ejector-ramjet. The performance characteristics of the IRS cycle are analyzed for a new single-stage-to-orbit (SSTO) launch vehicle concept, known as "Trailblazer." The study is based on a quasi-one-dimensional model of the rocket and air streams at speeds ranging from lift-off to Mach 3. The numerical formulation is described in detail. A performance comparison between the IRS and ejector-ramjet cycles is also presented.

Airborne mass spectrometers: four decades of atmospheric and space research at the Air Force research laboratory.

Abstract: Mass spectrometry is a versatile research tool that has proved to be extremely useful for exploring the fundamental nature of the earth's atmosphere and ionosphere and in helping to solve operational problems facing the Air Force and the Department of Defense. In the past 40 years, our research group at the Air Force Research Laboratory has flown quadrupole mass spectrometers of many designs on nearly 100 sounding rockets, nine satellites, three Space Shuttles and many missions of high-altitude research aircraft and balloons. We have also used our instruments in ground-based investigations of rocket and jet engine exhaust, combustion chemistry and microwave breakdown chemistry. This paper is a review of the instrumentation and techniques needed for space research, a summary of the results from many of the experiments, and an introduction to the broad field of atmospheric and space mass spectrometry in general.

Gas generating eject motor

Abstract: A gas generating eject motor (10) includes a case (12) containing an ignitable low temperature gas generant material (24) that does not produce toxic gases upon the combustion thereof. The gas generant material is generally contained with a screen enclosure (26, 30) housed within the case. An igniter (70) is disposed within the gas generant material for selectively igniting the gas generant to thereby generate combustion gases. A nozzle (40) is disposed within an open aft end of the case for focusing and directing the combustion gases generated by the ignited gas generant material. The case is constructed and arranged to be separaly attached to the aft end of a rocket (90) to be launched from a launch platform (80), so that, upon ignition of the gas generant, the combustion gases focused by the nozzle will apply a thrust to the rocket and thereby propel, or eject, the rocket from the launch platform, at which time the combustible propellant of the rocket motor will ignite and the eject motor will be separated from the rocket.

Effusion Cooling in Rocket Combustors Applying Fiber Reinforced Ceramics

Abstract: Since 1994 DLR has focused its efforts on combustion chamber technologies within the internal program "High Pressure Rocket Propulsion" (HDR). The main emphasis lays on advanced cooling technologies and especially effusion cooling applying fibre reinforced ceramics as porous media and hydrogen as cooling fluid. This paper summarises the recent experimental work on this cooling technique using Carbon/Carbon (C/C) material for combustion chamber components. After a brief summary of the fabrication process of the specific ceramic and the material properties, the basic equations for the flow situation and appropriate models are presented. Within the experimental campaign which has been performed at the DLR micro combustor facility M3 the combustion chamber pressure has been varied between 0.3 MPa and 1.1 MPa at a constant oxidiser/fuel mixture ratio of 6.5 using ambient hydrogen as coolant. The porosity of the ceramics has been varied in the range from about Epsilon = 13% to Epsilon = 25% in order to optimise the coolant mass flow rate and pressure loss across the porous wall. A wide range of coolant mass flow rates has been tested in order to achieve a broad data base for modelling an check the limits of applicability of the cooling technique and the ceramic wall material. The paper concludes with a presentation of experimental and theoretical results and a brief outlook to future activities.

Solid-fuel regression rate and combustion behavior of vortex hybrid rocket engines

Abstract: A series of static engine firings were conducted to investigate the solid-fuel regression rate behavior and operating characteristics of vortex hybrid rocket engines. The vortex hybrid engine configuration is characterized by a unique coaxial, co-swirling, counter-flowing vortex combustion field in the cylindrical fuel port. To generate this type of flow field, gaseous oxygen was injected through a swirl injector located between the aft-end of the fuel grain and the inlet to the converging portion of the exit nozzle. Test firings with thrusts up to 960 N (215 lb3 were conducted with GOX and HTPB. Fuel regression rates up to 650% larger than those in similar classical hybrids were measured. Empirical correlations were developed to accurately describe the experimental regression rates over more than an order of magnitude variation in mass flux. In addition to local mass flux and GOX injection velocity, geometric engine variables, such as engine contraction ratio and length-to-diameter ratio, had a significant influence on the measured regression rates. Physically descriptive, non-dimensional regression rate and heat transfer correlations were also developed. Throttling and re-start capability were demonstrated.

MASERATI: a RocketBorne tunable diode laser absorption spectrometer.

Abstract: The MASERATI (middle-atmosphere spectrometric experiment on rockets for analysis of trace-gas influences) instrument is, to our knowledge, the first rocket-borne tunable diode laser absorption spectrometer that was developed for in situ measurements of trace gases in the middle atmosphere. Infrared absorption spectroscopy with lead salt diode lasers is applied to measure water vapor and carbon dioxide in the altitude range from 50 to 90 km and 120 km, respectively. The laser beams are directed into an open multiple-pass absorption setup (total path length 31.7 m) that is mounted on top of a sounding rocket and that is directly exposed to ambient air. The two species are sampled alternately with a sampling time of 7.37 ms, each corresponding to an altitude resolution of approximately 15 m. Frequency-modulation and lock-in techniques are used to achieve high sensitivity. Tests in the laboratory have shown that the instrument is capable of detecting a very small relative absorbance of 10-4–10-5 when integrating spectra for 1 s. The instrument is designed and qualified to resist the mechanical stress occurring during the start of a sounding rocket and to be operational during the cruising phase of the flight when accelerations are very small. Two almost identical versions of the MASERATI instrument were built and were launched on sounding rockets from the Andoya Rocket Range (69 °N) in northern Norway on 12 October 1997 and on 31 January 1998. The good technical performance of the instruments during these flights has demonstrated that MASERATI is indeed a new suitable tool to perform rocket-borne in situ measurements in the upper atmosphere.

Elements of space technology for aerospace engineers

Abstract: Preface. Reference Frames and Time. Reference Frames. Motion in Accelerated Reference Frames. Example: The Yo-Yo Despin Mechanism. Euler Angles and Transformations of Coordinates. Time Intervals and Epoch. Forces and Moments. Gravity. Thrust. Aerodynamic Forces and Moments. Free Molecule Flow. Solar Radiation Pressure. Atmospheric Entry. Orbits and Trajectories in an Inverse Square Field. Kepler Orbits and Trajectories. Position as a Function of Time. D'Alembert and Fourier-Bessel Series. Orbital Elements. Spacecraft Visibility Above the Horizon. Satellite Observations and the f and g Series. Special Orbits. Perturbations by Other Astronomical Bodies. Planetary Fly-By and Gravity Assist. Relativistic Effects. Chemical Rocket Propulsion. Configurations of Liquid-Propellant Chemical Rocket Motors. Configurations of Solid-Propellant Motors. Rocket Stages. Idealized Model of Chemical Rocket Motors. Ideal Thrust. Rocket Motor Operation in the Atmosphere. Two and Three-Dimensional Effects. Critique of the Ideal Model. Elements of Chemical Kinetics. Chemical Kinetics Applications to Rocket Motors Liquid Propellants. Propellant Tanks. Propellant Feed systems of Launch Vehicles. Thrust Chambers of Liquid-Propellant Motors. Pogo Instability and Prevention. Thrust Vector Control. Engine Control and Operations. Liquid-Propellant Motors and Thrusters on Spacecraft. Components of Solid-Propellant Rocket Motors. Hybrid-Propellant Rocket Motors. Orbital Maneuvers. Minimum Energy Paths. Lambert's Theorem. Maneuvers with Impulsive Thrust. Hohmann Transfers. Other Transfer Trajectories. On-Orbit Drift. Launch Windows. Injection Errors and Their Corrections. On-Orbit Phase Changes. Rendez-Vous Maneuvers. Gravity Turn. Attitude Control. Principal Axes and Moments of Inertia of Spacecraft. The Euler Equations for Time-Dependent Moments of Inertia. The Torque-Free Spinning Body. Attitude Control Sensors. Attitude Control Actuators. Spin-Stablilized Vehicles. Gravity Gradient Stabilization. Spacecraft Thermal Design. Fundamentals of Thermal Radiation. Spacecraft Surface Materials. Model of a Spacecraft as an Isothermal Sphere. Earth Thermal Radiation and Albedo. Diurnal and Annual Variations of Solar Heating. Thermal Blankets. Thermal Conduction. Lumped Parameter Model of a Spacecraft. Thermal Control Devices. Bibliography. Appendix.

Cryogenic solid hybrid rocket engine and method of propelling a rocket

Abstract: A cryogenic solid hybrid engine with a solid propellant chamber, a first propellant within such chamber in which the first propellant is in solid form in the chamber and is in fluid form at room temperature, a coolant fluid chamber and a coolant fluid in the coolant fluid chamber being maintained at a temperature blow the freezing point of the first propellant. The invention also relates to a method for propelling a rocket and a method for forming a solid propellant grain for use in a cryogenic solid hybrid rocket engine.

Parachute device for helicopters

Abstract: A parachute device for a helicopter includes a canopy confining case having a stationary casing part, and a removable casing part mounted removably on the stationary casing part. The stationary and removable casing parts cooperatively form a compartment to receive a parachute canopy that has a release cord connected to the removable casing part, and a suspension line unit. An anchoring line has one end that extends into the canopy confining case and that is coupled to the suspension line unit of the parachute canopy. A rocket member includes a launch tube mounted on the stationary casing part externally of the compartment, and a rocket disposed in the launch tube and connected to the removable casing part. The rocket is capable of propelling from the launch tube when ignited. An ignition control line has a first end connected to the rocket member, and a second end provided with an ignition unit that is operable so as to ignite the rocket.