Showing papers on "Rocket published in 2008"

A Model of AP/HTPB Composite Propellant Combustion in Rocket-Motor Environments

Abstract: A comprehensive theoretical/numerical model for treating AP/HTPB composite-propellant combustion in a rocket-motor environment is presented. The formulation takes into account the conservation equations in both the gas and condensed phases, and accommodates finite-rate chemical kinetics and variable thermophysical properties. The processes in the two phases are coupled at the surface to determine the propellant burning behavior. An asymptotic analysis based on a large activation-energy approximation for the condensed-phase decomposition is applied to help resolve the combustion wave structure in the interfacial layer. A simplified global reaction is employed to characterize the final diffusion flame between the decomposition products of AP and the pyrolysis products of HTPB. Only laminar flows are considered here, to avoid complications arising from turbulence. A detailed parametric study is conducted on the gas-phase flame structures of AP/HTPB composite propellants. The dependence of burning rate, flame...

Computational Investigation of Acoustics and Instabilities in a Longitudinal-Mode Rocket Combustor

Abstract: A computational fluid dynamics analysis of acoustic modes and instabilities in an experimental longitudinal test chamber is presented. The experimental configuration is a uni-element recessed injector post combined with a variable-length combustion chamber. The computations employ the nonlinear Euler equations with mass and heat addition in the injector and combustion chamber and response functions to represent combustion dynamics. Analytical solutions and experimental comparisons are used to verify and validate the computational model. The results demonstrate the importance of including the full Euler equations for predicting the frequencies and mode shapes in the injector-combustor configuration as well as for representing nonlinear phenomena such as wave steepening and the excitation of higher harmonics. The present approach therefore promises to be a useful platform for testing and calibrating combustion response functions for combustion instability models.

Rocket and Spacecraft Propulsion

Abstract: History and principles of rocket propulsion -- The thermal rocket engine -- Liquid propellant rocket engines -- Solid propellant rocket motors -- Launch vehicle dynamics -- Electric propulsion -- Nuclear propulsion -- Advanced thermal rockets.The revised edition of this practical, hands-on book discusses the range of launch vehicles in use today throughout the world, and includes the very latest details of some of the advanced propulsion systems currently being developed. The author covers the fundamentals of the subject, from the basic principles of rocket propulsion and vehicle dynamics through the theory and practice of liquid and solid propellant motors, to new and future developments. The revised edition will stick to the same principle of providing a serious exposition of the principles and practice of rocket propulsion, but from the point of view of the user and enquirer who is not an engineering specialist. Most chapters will remain substantially the same as the first edition; they will be updated where necessary and errata corrected. The main revisions will be to the chapter on electric propulsion where there have been significant new developments both in engine types and in practical applications. This is now seen as the key to planetary exploration by robotic probes and should therefore be reflected. Nuclear propulsion has emerged from the doldrums and is now seen as a definite possibility for outer solar system robotic exploration; and as enabling technology for a human mars expedition. A new chapter on nuclear thermal propulsion has been added to reflect this revival of interest.

Test of B/Ti multilayer reactive igniters for a micro solid rocket array thruster☆

Abstract: In this study, reactive B/Ti multilayer igniters were investigated for the noncontact ignition of a micro solid rocket array thruster in vacuum. When current is supplied to the B/Ti multilayer igniter, the chemical reaction: 2B + Ti → TiB 2 + 1320 cal/g occurs, and sparkles are spread to a distance of several millimeters or more. The B/Ti multilayer igniters with three sizes were fabricated, and tested in six configurations of solid propellant. Although one rocket with ignition charge was ignited successfully, the noncontact ignition of the solid propellant was not achieved. However, the B/Ti multilayer igniters themselves generated small impulses of 10 −6 N s order, suggesting the possibility of self-propulsion.

Optimal Design of Hybrid Rocket Motors for Microgravity Platform

Abstract: A nested direct/indirect method is used to find the optimal design for a microgravity platform which is based on a hybrid sounding rocket. The direct optimization of the parameters that affect the motor design is coupled with the indirect trajectory optimization to maximize a given mission performance index. A gas-pressure feed system is used, with three different propellant combinations. The feed system exploits a pressurizing gas, namely, helium, when hydrogen peroxide or liquid oxygen is used as an oxidizer. The simplest blowdown design is compared with a more complex pressurizing system, which has an additional gas tank that allows for a phase with constant propellant tank pressure. Only self-pressurization is considered with nitrous oxide; two different models are used to describe the behavior of the tank pressurization. The simplest model assumes liquid/vapor equilibrium. A two-phase model is also proposed: Saturated vapor and superheated liquid are considered and the liquid/vapor mass transfer evaluation is based on the liquid spinodal line. Results show that the different tank-pressurization models yield minimal differences of the optimal motor characteristics. Performance differs slightly due to the different mass of the residual oxidizer. The propellant comparison for the present case shows better performance for hydrogen peroxide/ polyethylene with respect to liquid oxygen/hydroxyl-terminated polybutadiene, while nitrous oxide/hydroxyl-terminated polybutadiene remains attractive for system simplicity and low costs.

Lagrangian Trajectory Modeling of Lunar Dust Particles

Abstract: Apollo landing videos shot from inside the right LEM window, provide a quantitative measure of the characteristics and dynamics of the ejecta spray of lunar regolith particles beneath the Lander during the final 10 [m] or so of descent. Photogrammetry analysis gives an estimate of the thickness of the dust layer and angle of trajectory. In addition, Apollo landing video analysis divulges valuable information on the regolith ejecta interactions with lunar surface topography. For example, dense dust streaks are seen to originate at the outer rims of craters within a critical radius of the Lander during descent. The primary intent of this work was to develop a mathematical model and software implementation for the trajectory simulation of lunar dust particles acted on by gas jets originating from the nozzle of a lunar Lander, where the particle sizes typically range from 10 micron to 500 micron. The high temperature, supersonic jet of gas that is exhausted from a rocket engine can propel dust, soil, gravel, as well as small rocks to high velocities. The lunar vacuum allows ejected particles to travel great distances unimpeded, and in the case of smaller particles, escape velocities may be reached. The particle size distributions and kinetic energies of ejected particles can lead to damage to the landing spacecraft or to other hardware that has previously been deployed in the vicinity. Thus the primary motivation behind this work is to seek a better understanding for the purpose of modeling and predicting the behavior of regolith dust particle trajectories during powered rocket descent and ascent.

Thrust and Thermal Characteristics of Electromagnetic Launcher Based on Permanent Magnet Linear Synchronous Motors

Abstract: Compared with steam catapult system, electromagnetic launcher (EML) system is highly integrated, and it has high and well matching performance. It will be widely used aircraft carriers ejection, rocket launchers etc in future. Double-side tabular permanent magnet linear synchronous motor (PMLSM) for electromagnetic launcher can meet the requirements of big thrust and high efficiency etc. It can accelerate the launcher at the expected speed in short time. The thrust characteristic of the launcher is essential to the whole electromagnetic launcher system. Large thrust and small thrust ripple are both expected. This paper studies thrust characteristic of different pole arc coefficients, compares the series and parallel magnetic circuit structures, and analyses the method of staggering a certain distance between poles on both sides. In order to achieve the goal of large thrust, the launcher is often designed with high current density. As a result, it is of great loss and has quick temperature rise. This paper establishes numerical model of the two-dimensional (2D) primary transient temperature field. The temperature field of the short-term system is analyzed. The calculated and experimental results of the electromagnetic launchers can provide a basis for optimum design.

Correlation of Hybrid Rocket Propellant Regression Measurements with Enthalpy-Balance Model Predictions

Abstract: An enthalpy-balance fuel-grain regression model is presented. The regression model, based on the longitudinally averaged fuel recession rates, is shown to accurately predict the chamber pressure, thrust and specific impulse performance of small and medium scale hybrid rocket motors. The key to the model predictions is the longitudinal enthalpy balance between the fuel grain heat of ablation and the convective heat transfer from the flame zone to the model surface. Convective heat transfer is related to the surface skin friction using the Reynolds analogy for turbulent flow. Simple flat plate models are used to predict the longitudinally averaged skin friction coefficient. Chemical properties of the combustion products were evaluated using the NASA Computer Equilibrium with Applications (CEA) Combustion code. Model predictions for a nitrous oxide (N 2 O) and hydroxylterminated poly butadiene (HTPB) motor are compared to data from a small-scale test firing with a 10.2-cm diameter motor. Suggestions for model improvements are offered.

A Practical, Affordable Cryogenic Propellant Depot Based on ULA's Flight Experience

Abstract: Mankind is embarking on the next step in the journey of human exploration. We are returning to the moon and eventually moving to Mars and beyond. The current Exploration architecture seeks a balance between the need for a robust infrastructure on the lunar surface, and the performance limitations of Ares I and V. The ability to refuel or top-off propellant tanks from orbital propellant depots offers NASA the opportunity to cost effectively and reliably satisfy these opposing requirements. The ability to cache large orbital quantities of propellant is also an enabling capability for missions to Mars and beyond. This paper describes an option for a propellant depot that enables orbital refueling supporting Exploration, national security, science and other space endeavors. This proposed concept is launched using a single EELV medium class rocket and thus does not require any orbital assembly. The propellant depot provides cryogenic propellant storage that utilizes flight proven technologies augmented with technologies currently under development. The propellant depot system, propellant management, flight experience, and key technologies are also discussed. Options for refueling the propellant depot along with an overview of Exploration architecture impacts are also presented.

Apparatus For Protecting A Target From An Explosive Warhead

Abstract: A web includes strings and connectors that form ogive damagers. An ogive damager has three or more strings and three or more connectors. The connectors connect the strings to form a closed loop having an area that allows at least a tip of an ogive of a rocket to pass through the area. Each ogive damager is configured to damage the rest of the rocket.

Guidance Algorithm for Range Maximization and Time-of-Flight Control of a Guided Projectile

Abstract: The availability of gun-hardened guidance and control systems has made highly accurate gun-launched rocket-assisted guided projectiles feasible. A composite guidance algorithm is presented for such vehicles. The algorithm is capable of extending range and cross-range capability of the projectile, and allows it to be retargeted after launch. The algorithm also employs model predictive control to control time of flight to allow a salvo of projectiles to arrive simultaneously. The time-of-flight control achieves its objective by trajectory shaping and corrects for winds, off-nominal launch conditions, and rocket motor variations.

The ST5000: a high-precision star tracker and attitude determination system

Abstract: The University of Wisconsin's Space Astronomy Laboratory has designed and built a Star Tracker suitable for use on sounding rockets and class D satellites. This device brings together autonomous attitude determination ("Lost in Space" mode), multi-star tracking, and a novel form of Progressive Image Transmission (US patent #5,991,816), which allows the device to be used as an ultra-low bandwidth imager. The Star Tracker 5000 (ST5000) reached operational status in a suborbital sounding rocket flight in August 2007. The ST5000 determined the rocket's inertial (FK5) attitude with arcsecond precision using its autonomous attitude determination capability, and then provided continuous sub-arc-second tracking for the full 360-second on-target portion of the flight. The ST5000 RMS tracking error was 0.54 arc-seconds in yaw and pitch, and 17 arc-seconds in roll. The vehicle RMS jitter was 0.5 arc-seconds in yaw and pitch, and 10 arc-seconds in roll. The ST5000 was funded by NASA grants NAG5-7026 and NAG5-8588.

A Contribution to the Development of Actual Continuous Detonation Wave Engine

Abstract: During past years, MBDA performed some theoretical and experimental works, mainly in cooperation with LCD laboratory at ENSMA Poitiers, on Pulsed Detonation Engine (PDE) These studies aimed at obtaining a preliminary demonstration of the feasibility of the PDE in both rocket and airbreathing modes and at verifying the interest of such a PDE for operational application : rocket and airbreathing mode experimental evaluation, effect of filling coefficient, effect of a nozzle, thermal, mechanical, acoustic and vibrations environment generated, evaluation of different fuels, performance code development Due to its thermodynamic cycle, the pulsed detonation engine (PDE) has theoretically a higher performance than an other classical propulsion concept using the combustion process (+ 20 to 25% in term of thermal efficiency) Nevertheless, it is necessary to verify that this advantage is not fully compensated by the difficulties, which could be encountered for practical use of the PDE concept or by the complex technology, which could be needed to implement it in an operational flying system Moreover, a PDE a priori generates a severe vibration environment, which can imply higher more severe requirement for all on-board vehicle equipments or subsystems After some in house studies performed in national and international cooperation, MBDA is now focusing its efforts to the development of a demonstration engine led in cooperation with Singapourian DSO in order to really assess the feasibility and the interest of this propulsion concept

Solid Rocket Combustion Instability—An American Historical Account

Abstract: T combustion processes are both important in rocket motors and difficult to understand. A wide range of transient processes are of interest, including ignition, quenching, transition from deflagration to detonation, and oscillatory combustion. These problems are often determined partly by factors outside the combustion zone of the propellant (e.g., an igniter, gas oscillations, etc.), but they all involve transient behavior of the combustion process, which is the central theme of this book. Much of the research on steady and transient combustion has been motivated by the problem of combustor instability in rocket motors that sometimes leads to destructive combustion-driven gas oscillations in the combustor. This chapter is a historical review of progress in the oscillatory combustor instability problem. Days of the Magicians

Handling Considerations of Nitrous Oxide in Hybrid Rocket Motor Testing

Abstract: Nitrous Oxide has many beneficial properties that make it a good choice for use in hybrid rocket motors. The large quantities and potentially high pressures used in rocket motors present unique hazards that are not generally found in the standard industrial and medical uses of nitrous oxide. Through years of hands on use, and research into the properties of nitrous oxide, SpaceDev has created a set of guidelines on how to design, clean, and inspect systems using nitrous oxide. This paper presents the nitrous oxide handling precautions and procedures employed by SpaceDev in the testing of hybrid rocket motors.

Opportunities for a Liquid Rocket Feed System Based on Electric Pumps

Abstract: A feed system for liquid-propellant rocket engines based on electric pumps powered by batteries is proposed. It is proven to stand as a viable alternative to the pressure-gas feed system. The dependence of the feed system mass on the different operating parameters is obtained so as to identify the conditions favoring its adoption, that is, a relatively long burning time and a fairly high chamber pressure. Under such conditions, the proposed system is shown to offer significant mass savings with respect to the pressure-gas system when advanced batteries are used. This advantage is further enhanced by the beneficial effect of chamber pressure on the engine effective exhaust velocity. A test case for a low Earth orbit to geostationary equatorial orbit transfer is also presented to identify the optimum value of the burning time, deriving from the competition between the feed system mass and the effect of gravitational losses.

Real gas CFD simulation of supercritical H2-LOX combustion in the Mascotte single-injector combustor using a commercial CFD code

Abstract: Modern high performance rocket combustion engines operate at very high pressures up to 20 MPa. Propellants, typically hydrogen and oxygen, are injected at very low temperatures, below the critical temperature of oxygen (154.581 K), whereas the pressure has a supercritical value (p > pc = 50.43 bar) so that mixing and combustion occur at transcritical conditions. The material properties and equation of state in this region differ significantly from those at low pressures and high temperatures. The paper presents an implementation of real gas models into the commercial CFD-Code ANSYS CFX. This solver is validated with experimental data measured on the Mascotte test rig (V03) at ONERA which have been published in the 2 nd International Workshop on Rocket Combustion Modelling.

Coupled analysis of flow and surface ablation in carbon-carbon rocket nozzles

Abstract: A study is conducted to predict C/C nozzle recession behavior in solid rocket motors for broad variations of propellant formulations and motor operating conditions. The numerical model considers the turbulent flow in the nozzle, heterogeneous chemical reactions at the nozzle surface, variable transport and thermodynamic properties, and heat conduction in the nozzle material. Results show that the recession rate is largely determined by the diffusion of the major oxidizing species (H2O, CO2, OH) to the nozzle surface. Both the concentration of the major oxidizing species -affected by the aluminum content of the propellant- and the chamber pressure exert a strong influence on the recession rate. The erosion rate increases almost linearly with chamber pressure and decreases with propellants with higher aluminum content. The calculated results show a very good agreement with the experimental data from the BATES motor firings.

Optimal singular rocket and aircraft trajectories

Abstract: We give a brief review of optimal singular control rocket and aircraft trajectories in outer space and in the atmosphere. Singular control rocket and aircraft trajectories occur when the thrust is at intermediate levels. The analysis of a singular control rocket or aircraft trajectory requires the Goh-Legendre-Clebsch necessary conditions. Two sets of compact forms of these conditions are described. The first compact form of these optimality conditions is convenient for testing the optimality of a singular control trajectory. The second compact form is important because its strengthened form provides sufficient conditions for the control variables to be expressed in terms of the state and costate variables.

The QED engine system: Direct‐electric fusion‐powered rocket propulsion systems

Abstract: Practical ground‐to‐orbit and inter‐orbital space flights both require propulsion systems of large flight‐path‐averaged specific impulse (Isp) and engine system thrust‐to‐mass‐ratio (F/me=[F]) for useful payload and structure fractions in single‐stage vehicles (Hunter 1966). Current rocket and air‐breathing engine technologies lead to enormous vehicles and small payloads; a natural result of the limited specific energy available from chemical reactions. While nuclear energy far exceeds these specific energy limits (Bussard and DeLauer 1958), the inherent high‐Isp advantages of fission propulsion concepts for space and air‐breathing flight (Bussard and DeLauer 1965) are negated for manned systems by the massive radiation shielding required by their high radiation output (Bussard 1971). However, there are well‐known radiation‐free nuclear fusion reactions (Gross 1984) between isotopes of selected light elements (such as H+11B, D+3He) that yield only energetic charged particles, whose energy can be converted directly into electricity by confining electric fields (Moir and Barr 1973,1983). New confinement concepts using magnetic‐electric‐potentials (Bussard 1989a) or inertial‐collisional‐compression (ICC) (Bussard 1990) have been found that offer the prospect of clean, compact fusion systems with very high output and low mass. Their radiation‐free d.c. electrical output can power unique new electron‐beam‐driven thrust systems of extremely high performance. Parametric design studies show that such charged‐particle electric‐discharge engines (‘‘QED’’ engines) might yield rocket propulsion systems with performance in the ranges of 2<[F]<6 and 1500

Thermonuclear plasma reactor for rocket thrust and electrical generation

Abstract: A reactor system produces plasma rocket thrust using alpha-initiated atomic fuel pellets without the need for a critical mass of fissionable material. The fuel pellets include an outer layer reactive material to alpha particles to generate neutrons (e.g., porous lead or beryllium), an under-layer of fissionable material (e.g., thorium or enriched uranium), and an optional inner core of fusion material (e.g., heavy water ice, boron hydride). The pellets are injected one at a time into a charged reaction chamber containing a set of alpha beam channels, possibly doubling as ion accelerators, all directed toward a common point. Alpha particles converging on each successive pellet initiate an atomic reaction in the fissionable under-layer, via a neutron cascade from the pellet outer layer, producing plasma that is confined within the chamber. This may be enhanced by atomic fusion of the optional inner core. The resulting high-energy plasma creates electrostatic pressure on the chamber and is allowed to exit the chamber through a port. An ion accelerator at the exhaust port of the chamber accelerates outgoing plasma ions, possibly with added reaction mass, to generate the rocket thrust. An electric circuit that includes the charged chamber may collect the electrons in the plasma to help power the ion accelerator(s).

Multifunctional artificial hail-proof rain-increasing rocket pod and emission controller thereof

Abstract: A multifunctional artificial anti-hail rocket launcher has at least two groups of rocket launch trajectory assemblies which are fixed on a trajectory support assembly. An upper base is mounted in a lower base and can turn. The bottom of the trajectory support is articulated with the upper base, between which and the trajectory support is mounted a pinching stretching support assembly. A mating launch controller of the multifunctional artificial anti-hail rocket launcher includes a comparison and display circuit, a test/launch transformation circuit, launch circuit and a power supply. The utility model is structurally simple and concise, widely applicable and easy to use. Not only the position but also the pinching angle can be adjusted expediently. The utility model can be applied to rocket launchings of different specifications and factories. Just a single set can be applied to artificial weather modification in different weather conditions, object cloud systems and altitudes, thus reducing the artificial weather modification cost greatly.

A CFD-RTE Model for Thermal Analysis of Regeneratively Cooled Rocket Engines

Abstract: This paper presents a computational methodology for thermal analysis of hot-gas-side and coolant–side of regeneratively cooled liquid rocket engines. The computational methodology is composed of a CFD model, for the hot-gas-side and RTE (Rocket Thermal Evaluation) for the coolant flow and wall conduction. The CFD model solves the axisymmetric flow and thermal fields of hot-gas in the thrust chamber and nozzle. The RTE predicts the coolant side properties and the wall temperature distribution. This integrated CFD-RTE model is validated by comparing its results with the published data for the space shuttle main engine (SSME). Nomenclature A r = surface area vector f A r = area of face f faces N = number of faces enclosing cell φ S = source of φ per unit volume V = cell volume φ

Ram and Ejector-Jet Modes Experiments of the Combined Cycle Engine in Mach 4 Flight Conditions

Abstract: A combustion-capable combined cycle engine model which was constructed based on the rocket and ramjet technology was tested in Mach 4 flight condition. At this speed, engine is designed to shift its operation mode from an ejector-jet to a ramjet. Both modes were simulated by changing the rocket combustion pressure. Even with full rocket exhaust, no effect to the air flow could be observed. The injection point of the secondary fuel affected thrust performance. In the ramjet mode, the pressure rise due to the fuel combustion traveled to the entrance of the combustor, but it stayed near the injection point in the ejector-jet mode.

Overview of rocket engine control

Abstract: This paper broadly covers the issues of Chemical Rocket Engine Control. The basic feedback information and control variables used in expendable and reusable rocket engines, such as the Space Shuttle Main Engine are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control systems for rocket engines (and vehicles) is presented.

Electromagnetic launch assistance for space vehicles

Abstract: Motivated by reducing the cost of launching space vehicles, the US National Aeronautics and Space Administration sponsored an industry/university research project to explore the application of electromagnetic forces. Here, the technical issues involved are examined, the initial achievements reported and further improvements suggested. If the initial launch phase of a space vehicle is horizontal instead of vertical, it is possible to use electromagnetic forces to supplement the rocket thrust, with consequent saving of rocket fuel. The short time of the horizontal launch phase (10-20s) enables induction levitation to be used in combination with a compact form of a linear induction motor. The performance, dynamic characteristics and scaling laws of this system are examined. Induction levitation is simple and effective, but it has an oscillatory response to force or torque disturbances, and it is unsuitable for very large space vehicles. These problems can be overcome with magnetic suspension using controlled DC electromagnets. Although the energy required for electromagnetic launch assistance is small, the electrical power demand is very high, necessitating some form of local energy storage. The Institution of Engineering and Technology 2008

A Comparison of Properties of Aluminized Composite Propellants Containing HMX and FOX-7

Abstract: Modern rocket propellants contain inter alia nitroamines (i.e. RDX, HMX). Therefore, the detonation properties of composite solid propellants are very important for good functioning of rocket motors and for storage. One of the new materials of that kind, with low sensitivity, is FOX-7 which was applied here as one of the components of composite solid propellant. This paper presents the thermodynamical calculations and thermochemical research results as well as the results of a study of transition into detonation of an aluminized composite propellant containing HMX. The said properties were compared with those of a propellant containing FOX-7.