Showing papers on "Rocket published in 2010"

Liquid-propellant rocket engine throttling: A comprehensive review

Abstract: Liquid-propellant rocket engines are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable liquid-propellant rocket engines can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable liquid-propellant rocket engines can also continuously follow the most economical thrust curve in a given situation, as opposed to making discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an liquid-propellant rocket engine as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of liquid-propellant rocket engine throttling centered around engines from the United States. Several liquid-propellant rocket engine throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

A new Langmuir probe concept for rapid sampling of space plasma electron density

Abstract: In this paper we describe a new Langmuir probe concept that was invented for the in situ investigation of HF radar backscatter irregularities, with the capability to measure absolute electron density at a resolution sufficient to resolve the finest conceivable structure in an ionospheric plasma. The instrument consists of two or more fixed-bias cylindrical Langmuir probes whose radius is small compared to the Debye length. With this configuration, it is possible to acquire absolute electron density measurements independent of electron temperature and rocket/satellite potential. The system was flown on the ICI-2 sounding rocket to investigate the plasma irregularities which cause HF backscatter. It had a sampling rate of more than 5 kHz and successfully measured structures down to the scale of one electron gyro radius. The system can easily be adapted for any ionospheric rocket or satellite, and provides high-quality measurements of electron density at any desired resolution.

Operability Benefits of Airbreathing Hypersonic Propulsion for Flexible Access to Space

Abstract: Hypersonic airbreathing propulsion has been considered as an enhancement for access-to-space systems for decades. However, previous research usingmetrics such as takeoff gross weight and payload weight fraction has not shown conclusive benefits for airbreathing systems when compared with all-rocket launch vehicles. The U.S. Air Force Research Laboratory has developed new operability-based metrics relevant to U.S. Air Force missions: time to rendezvous with a target spacecraft, number of launch opportunities per day, and launch-window duration. Computation of the new metrics requires launch vehicle ascent trajectory optimization, orbital transfer solutions, andMonteCarlo analysis. Ascent optimization uses propulsion throttling, aerodynamic turning, andpitch control to command downrange and crossrange at the orbital insertion point while using minimum propellant. Then the twopoint boundary-value problem is solved to find a minimum-propellant transfer orbit to rendezvous with the target. Monte Carlo analysis assigns the orbital target a random starting position over the Earth and then propagates the orbit until rendezvous is accomplished and themetrics can be computed. The Air Force’s ReusableMilitary Launch System all-rocket launch vehicle RMLS 102 is compared against Alliant Techsystems’ rocket-based combined-cycle launch system ATK-RBCC. The results of the Monte Carlo analysis show that ATK-RBCC featuring airbreathing propulsion combined with a high lift-over-drag airframe exhibits significant operability benefits. The developed operability metrics could help to transform access to space by demonstrating clear payoffs from airbreathing propulsion.

Numerical Investigation of the Effect of a Diaphragm on the Performance of a Hybrid Rocket Motor

Abstract: This paper describes the use of a CFD code (Ansys CFX 12) for the analysis of a hybrid rocket motor with a diaphragm placed in the combustion chamber in order to enhance rocket performance. This work follows the experimental campaign of Matthias Grosse who tested the motor using nitrous oxide and paraffin wax as propellants. Several of his tests have been used as a reference for the numerical simulations. Several approximations have been made: steady state conditions, eddy dissipation combustion model with one-step reaction, gaseous injection of fuel and oxidizer, no droplets entrainment (typical of a paraffin grain). First of all, a single geometry without diaphragm has been analyzed with different turbulence models (k-ω, k-ω SST, k-e, k-e RNG). It has been shown that the k-ω model predicts a lower flame temperature and chamber pressure than k-e. Five geometries have been studied in order to compare the use of two different types of diaphragm (1 hole and 4 holes) in two positions (24% and 33% of the total length) respect to a configuration without mixer. The effect of the diaphragm is an increase of the mixing of the chemical species participating in the combustion process. The use of the diaphragm showed a performance enhancement, as showed in the experimental campaign. There is a good agreement between CFD results and experimental data: the efficiency is overestimated by less than 5.5%. This work proves the capability of CFD codes to predict global hybrid motor performances and to be a useful tool in the study of mixing devices.

Detonation Limit Thresholds in H2/O2 Rotating Detonation Engine

Abstract: The rotating detonation engine (RDE) is a new engine system using detonation, which may provide a higher performance and smaller and simpler design in comparison with the pulse detonation engine (PDE) and other traditional engines. However the research on RDE stands just at the first step now. The authors perform a numerical analysis to understand about RDE in terms of features of rotating detonation and its propagation limit. The lower threshold pressure of detonation limit was 2.6 MPa and the upper threshold pressure of detonation limit was 7.1 MPa. The engine performance analysis shows that the maximum mixture based specific impulse (I spm ) was about 440 s, which is comparable with that of the present typical rocket engine.

Optimal Design of Hybrid Rocket Motors for Launchers Upper Stages

Abstract: A hybrid rocket is considered as the third stage of a three-stage launcher. The propulsion system design and the trajectory are simultaneously optimized by means of a nested direct/indirect procedure. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload for assigned conditions at the stage ignition and final orbit. A mission profile based on the Vega launcher is considered. The feed system exploits a pressurizing gas, namely helium, with hydrogen peroxide as the oxidizer and polyethylene as the fuel. 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 oxidizer tank pressure. The optimization provides the optimal values of the main engine design parameters (pressurizing gas mass, nozzle expansion ratio, and initial values of tank pressure, mixture ratio and thrust), the corresponding grain and engine geometry, and the control law (thrust direction during the ascent trajectory and engine switching times). Results show that a hybrid rocket may be a viable option for small launchers.

Film Cooling in a High-Pressure Subscale Combustion Chamber

Abstract: By the application of film cooling in addition to regenerative cooling, a considerable reduction in thermal and structural loads of rocket combustion-chamber walls can be reached. This paper discusses important influence parameters on film cooling in terms of efficiency of the injected film and wall temperature reduction. For the experimental investigations a high-pressure subscale combustion chamber operated with the cryogenic propellant combination LOX=GH2 was used. A gaseous film with ambient tempered hydrogen was injected in the axial direction at the face plate. Typical film-cooling parameters such as film blowing rate, velocity ratio between film injection velocity and hot-gas velocity, circumferential slot positioning, and film injection slot height were investigated systematically at the European Research and Technology Test Facility P8.

Investigation of Rocket Exhaust Diffusers for Altitude Simulation

Abstract: The design and operational parameters of rocket exhaust diffusers equipped to simulate high-altitude rocket performance on the ground were investigated and characterized using a comprehensive approach (theoretical, numerical, and experimental). The physical model of concern includes a rocket motor, a vacuum chamber, and a diffuser, which have axisymmetric configurations. Further, the operational characteristics of a rocket exhaust diffuserwereanalyzed froma flowdevelopmentpointof view.Emphasiswasplacedondetailed flowstructure inthe diffuser, to observe the pressure oscillation in both the vacuum chamber and diffuser, which determines the minimum rocket-motor pressure required to start the diffuser. Numerical simulations were compared with experimental data on startup and in operational conditions to understand the effects of major design parameters, including the area ratio of diffuser to rocket-motor nozzle throat, the rocket-motor pressure, and the vacuumchamber size. Nomenclature Ad = inner cross-sectional area of diffuser Ade = exit cross-sectional area of diffuser Ae = exit cross-sectional area of rocket nozzle At = throat cross-sectional area of rocket nozzle

Control system design and experiment of needle-type gas regulating system for ducted rocket:

Abstract: The control system design and experimental investigation of the gas regulating system for a ducted rocket are studied in this article. First, the dynamic model of the gas regulating system ...

Scaling of Erosion Rate in Subsonic Jet Experiments and Apollo Lunar Module Landings

Abstract: Small scale jet-induced erosion experiments are useful for identifying the scaling of erosion with respect to the various physical parameters (gravity, grain size, gas velocity, gas density, grain density, etc.), and because they provide a data set for benchmarking numerical flow codes. We have performed experiments varying the physical parameters listed above (e.g., gravity was varied in reduced gravity aircraft flights). In all these experiments, a subsonic jet of gas impinges vertically on a bed of sand or lunar soil simulant forming a localized scour hole beneath the jet. Videography captures the erosion and scour hole formation processes, and analysis of these videos post-test identifies the scaling of these processes. This has produced important new insights into the physics of erosion. Based on these insights, we have developed an erosion rate model that can be applied to generalized situations, such as the erosion of soil beneath a horizontal gas flow on a planetary surface. This is important to lunar exploration because the rate of erosion beneath the rocket exhaust plume of a landing spacecraft will determine the amount of sand-blasting damage that can be inflicted upon surrounding hardware. Although the rocket exhaust plume at the exit of the nozzle is supersonic, the boundary layer on the lunar surface where erosion occurs is subsonic. The model has been benchmarked through comparison with the Apollo landing videos, which show the blowing lunar soil, and computational fluid dynamics simulations of those landings.

Frequency-Dependent FDTD Simulation of the Interaction of Microwaves With Rocket-Plume

Abstract: The ionized exhaust plumes of solid rocket motors may interfere with RF transmission under certain flight conditions. To understand the important physical processes involved, we measured microwave attenuation and phase delay due to the exhaust plume during sea-level static firing tests for a full-scale solid propellant rocket motor. The measured data were compared with the results of a detailed simulation performed using the frequency-dependent finite-difference time-domain ((FD)2TD) method. The numerically derived microwave attenuation was in good agreement with experimental data. The results revealed that either the line-of-sight microwave transmission through ionized plumes or the diffracted path around the exhaust plume mainly affects the received RF level, which depends on the magnitude of the plasma-RF interaction.

High Speed Visualization of LOX/GH2 Rocket Injector Flowfield: Hot-Fire and Cold-Flow Experiments

Abstract: The primary atomization and combustion characteristics of a liquid oxygen (LOX) / gaseous hydrogen (GH2) shear coaxial injector element were experimentally investigated. High speed movies using a shadowgraph imaging technique to visualize the LOX core were recorded for both hot-fire (LOX/GH2) and cold-flow (LOX/gaseous oxygen (GO2)) conditions with the same injector and chamber. Flow conditions were set to approximate realistic rocket conditions. For the hot-fire tests (LOX/GH2), chamber pressures were 600, 730, and 920 psia, with momentum flux ratios (annulus flow/post flow) of 2.7, 2.0 and 1.6 respectively. The rocket assembly utilized a preburner to provide a background flow (M≈0.1) of hot gaseous nitrogen (GN2)/GH2/water (H2O) gas with 25% volumetric concentration of hydrogen. For the cold-flow tests (LOX/GO2 with GO2 background flow), chamber pressures were 650 and 830 psia, thus above and below the critical pressure of oxygen (731.6 psia), with momentum flux ratios (annulus flow/post flow) of 2.2 and 1.8 respectively. The high speed visualizations under hot-fire conditions show a long sinuous LOX core region that breaks into large dense-oxygen structures, which are then quickly consumed. These results do not agree with the classical phenomenological breakup model that suggests a liquid core that is rapidly sheared into a drop cloud. Rather, a large-scale fragmentation model may be better suited to describe the primary atomization behavior in combusting flow from a LOX/GH2 shear coaxial injector element at realistic rocket conditions. Unlike the hot-fire case, cold-flow LOX visualization movies show a clear difference between the two chamber pressures, with the higher pressure (supercritical) case resembling behavior indicative of gaseous mixing compared to the typically two phase mixing appearance of the lower pressure (subcritical) case. Time-resolved measurements of the intact-core length are presented, along with size and frequency distributions of separating large dense-oxygen structures under hot-fire conditions.

A Reusable, Rocket and Airbreathing Combined Cycle Hypersonic Vehicle Design for Access-to-Space

Abstract: Several rocket boosted RBCC (Rocket-Based Combined Cycle) designs were created and analyzed for suitability for an access-to-space TSTO vehicle system. Level 1 analysis was performed that included vehicle closure, weight breakdown (to component level), flight trajectory data, propulsion and aerodynamic performance. Various inlet shapes were considered, such as symmetric and non-symmetric inlets. Also, several geometric configurations were studied, such as single flowpath, dual flowpath, engine-on-top and engine-on-bottom. At the end of the study, the number of candidate designs were reduced to two; one as a primary design and the other as the backup design. The primary design, a dual flowpath, engine-on-top design was selected for further analysis. Due to the large volume of the payload, the dual flowpath design was found to be more suitable than a single flowpath design.

Aerospace Propulsion Systems

Abstract: Preface Acknowledgments Outline of Aerospace Propulsion History Nomenclature 1 Fundamentals 11 Fundamental Equations 111 Review of Terms 112 Equation of State for a Perfect Gas 113 Law of the Conservation of Mass 114 Law of the Conservation of Linear Momentum 115 Law of the Conservation of Energy 12 Isentropic Equations 121 Isentropic Relationship between Temperature and Pressure 122 Isentropic Relationships with Specific Volume 13 Polytropic Processes 14 Total (or Stagnation) Properties 15 Isentropic Principles in Engine Components 151 Ducts 152 Turbomachinery 153 Combustion Chambers (Combustors) 154 Nozzles 16 Shock Waves 161 Normal Shocks 162 Oblique Shocks 163 Conical Shocks 17 Summary References Problems 2 Rockets 21 Background Description 22 Performance of an Ideal Rocket 221 Rocket Thrust Equation 222 Total and Specific Impulse 223 Effective Exhaust Velocity 224 Rocket Efficiencies 225 Characteristic Velocity 226 Thrust Coefficient 23 Solid Rocket Motors 231 Colloidal (Homogenous) Propellants 232 Composite (Heterogeneous) Propellants 233 Composite Modified Double-Based Propellants 234 Solid Propellant Grain Geometry 235 Solid Rocket Motor Casing 236 Combustion of Solid Propellants 237 Solid Rocket Ignition Systems 24 Liquid Rockets 241 Liquid Rocket Propellants 242 Liquid Rocket Feed Systems 243 Liquid Rocket Injection Systems 244 Combustion of Liquid Propellants 245 Liquid Rocket Ignition Systems 25 Hybrid Rockets 26 Motor Casing 27 Thrust Chamber 28 Exhaust Nozzles 29 Multi-staging 210 Non-chemical Rockets 211 Rocket Design Methodology 212 Summary References Problems 3 Piston Aerodynamic Engines 31 Background Description 32 Engine Types 321 Rotary Engines 322 Reciprocating Engines 323 Supercharged Reciprocating Engines 324 Gas Turbine Propeller Engines 33 Thrust 34 Combustion 341 Aviation Fuel 342 Specific Fuel Consumption 35 Propeller Design 351 General Description 352 Power Efficiencies 353 Variable Pitch Blades 354 Propeller Shapes 355 Contra-Rotating Propellers 356 Helicopter Rotor Blades 36 Propeller Performance 37 Summary References Problems 4 Gas Turbine Engines 41 Background Description 42 Ideal Gas Turbine Cycle 43 Types of Gas Turbine Engines 431 Jet Propulsion Engines 432 Shaft Power Engines 44 Engine Cycle Performance 441 Jet Propulsion Thrust 442 Shaft Power Thrust 443 Propulsive Efficiency 444 Thermal Efficiency 445 Overall Efficiency 446 Specific Fuel Consumption 45 Component Performance 451 Intakes 452 Compressors 453 Turbines 454 Combustion Chambers (Combustors) 455 Exhaust Nozzles 46 Engine Performance Analysis 47 Design Point Optimization 48 Component Design 481 Intake Design 482 Compressor System Design 483 Combustion Chambers 484 Turbines 485 Exhaust Systems 49 Engine Control Systems 410 Summary References Problems 5 Ramjet and Scramjet Engines 51 Background Description 52 Ramjet Engines 521 Conventional Ramjet Engines 522 Turboramjet Engines 523 Analysis of Ramjet Engines 524 Ramjet Component Design 53 Scramjet Engines 531 Description 532 Scramjet Component Design 54 Summary References Problems Appendix A: Gas Tables Appendix B: Isentropic Flow Tables Appendix C: Shock Tables Appendix D: Rocket Propellant Tables Solutions to Even Numbered Problems Index

Numerical calculation of FSS/RSS transition in highly overexpanded rocket nozzle flows

Abstract: Turbulent flow separation in over-expanded rocket nozzles is investigated numerically in a sub-scale parabolic nozzle fed with cold nitrogen. Depending upon the feeding to ambient pressure ratio either a free shock separation or a restricted shock separation is computed, with a significant hysteresis between these two flow regimes. This hysteresis was also found in experimental tests with the same nozzle geometry. The present study is mainly focused on the transition between the two shock separation patterns. The analysis of the numerical solutions aims to provide clues for the explanation of the hysteresis cycle.

Numerical Investigation of Supercritical Coolant Flow in Liquid Rocket Engine

Abstract: Understanding and predicting the flowfield and heat transfer characteristics in cooling channels are prerequisite to improve design and performance of regeneratively cooled rocket thrust chambers. In order to realize them, a CFD code, able to predict such characteristics, is developed based on a pressure-based solver and the cubic-type equation of state to take into account the real gas effect. As a preliminary study, simulations of transcritical parahydrogen flows in a uniformely heated circular tube are performed in order to validate the developed code against the reference experiment and investigate the flowfield and the heat transfer characteristics under transcritical conditions. The computed results agree well with the experimental data with regard to the wall temperature, the heat transfer coefficient, the bulk pressure and temperature. Also, the peculiar behavior, called “heat transfer deterioration”, under transcritical condition with high heat flux, is successfully predicted. The simulated flowfield reveal the mechanism of it. The parametric studies with different heat flux levels clarify the condition in which the heat transfer deterioration takes places.

A more thorough analysis of water rockets: Moist adiabats, transient flows, and inertial forces in a soda bottle

Abstract: The water rocket 1 is a popular toy that is often used in first year physics courses to illustrate Newton’s laws of motion and rocket propulsion. In its simplest version, a water rocket is made of a soda bottle, a bicycle pump, a rubber stopper, and some piping see Fig. 1. The bottle is half-filled with water, turned upside-down, and air is pushed inside the bottle via a flexible pipe that runs through the stopper. When the pressure builds up, the stopper eventually pops out of the neck. The water is then ejected and the rocket takes off. Witnesses of the launch of a water rocket cannot but be amazed that such a simple device can reach a height of tens of meters in a fraction of a second. The popularity of water rockets extends beyond physics classrooms, with many existing associations and competitions organized worldwide. 1 The more than 5000 videos posted on YouTube with the words “water rocket” in their title testify to their popularity. Some of these videos involve elaborate technical developments such as multistage water rockets, nozzles that adapt to the pressure, the replacement of water by foam or flour, underwater rocket launches, and even a water-propelled human flight. The public’s passionate explorations with water rockets contrast with the small number of articles devoted to their analysis. I found only two papers 2,3 that treat the simplest possible rocket, similar to

Nonlinear Liquid Rocket Combustion Instability Behavior using UCDS™ Process

Abstract: The Universal Combustion Device Stability™ (UCDS™) Process is the culmination of more than 40 years of research and provides the means to understand the complex dynamics and processes inside any chemical propulsion system, including rockets, turbojets and scramjets. With UCDS, it is now possible to accurately predict the time history of the pressure oscillation amplitudes during operation. This paper will provide an overview of the UCDS process as it is applied to liquid rocket engines. Using the UCDS nonlinear models a series of simulated bomb tests are used to explore the range of dynamic responses for a generic liquid rocket engine. These results reveal a wide range of complex dynamic responses that explain the widely varying instability effects seen in actual liquid rocket tests. When applied to the Corporal liquid rocket, this case study shows that the oscillatory behavior seen in the extensive testing of the Corporal rocket performed by JPL can be matched and clearly explained.

Photon rockets moving arbitrarily in any dimension

Abstract: A family of explicit exact solutions of Einstein's equations in four and higher dimensions is studied which describes photon rockets accelerating due to an anisotropic emission of photons. It is possible to prescribe an arbitrary motion, so that the acceleration of the rocket need not be uniform - both its magnitude and direction may vary with time. Except at location of the point-like rocket the spacetimes have no curvature singularities, and topological defects like cosmic strings are also absent. Any value of a cosmological constant is allowed. We investigate some particular examples of motion, namely a straight flight and a circular trajectory, and we derive the corresponding radiation patterns and the mass loss of the rockets. We also demonstrate the absence of "gravitational aberration" in such spacetimes. This interesting member of the higher-dimensional Robinson-Trautman class of pure radiation spacetimes of algebraic type D generalises the class of Kinnersley's solutions that has long been known in four-dimensional general relativity.

Simulation Code for Hybrid Rocket Combustion

Abstract: The simulation of the flame structure typical of hybrid rockets is addressed in this paper. The process is a complex interaction among oxidizer atomization and vaporization, gas phase combustion, fuel surface pyrolysis, soot formation and radiation phenomena. A detailed knowledge is required for the design of hybrid rocket motors and numerical simulations can be useful to improve grounding knowledge and reduce testing efforts. A hybrid flame consists of a turbulent reactive boundary layer with cross-flow blowing of gaseous fuel from the solid grain. Heat feedback from the flame sustains the combustion, being responsible for the pyrolysis of the solid fuel. The paper refers to a computational code for the simulation of hybrid flame structure and the prediction of fuel regression rate under development at the Space Propulsion Laboratory of Politecnico di Milano. The code uses a finite volume method and solves for Navier Stokes equations with RANS approach and specific combustion kinetic set. Closure terms are derived from Launder Sharma turbulence model and well stirred reactor model. Radiation is resolved through a P1 model. The computational domain is split into a solid and a gas phase region. The paper presents some typical results of the code for polybutadiene and gaseous oxygen as well as preliminary outcomes of a newly implemented two-phase flow model that simulates the combustion of aluminum drops in the core flow. Aluminum agglomerates are released in liquid state from the solid fuel surface and burn while being transported by the hot gases. The evolution of aluminum drops is derived from the Beckstead’s law. The simulation code is built on top of the open source framework OpenFOAM. A multi-domain approach is used simulating the reacting gas mixture on one side and the heat conduction in the solid fuel grain on the other.

CFD-Simulation of the injection and combustion of LOX and H2 at supercritical pressures

Abstract: Due to the high pressures and very low injection temperatures of the propellants in modern rocket combustors real gas effects play an important role in rocket combustion simulation. These have to be accurately modelled in combustion CFD simulations to enable a reliable prediction of the performance and lifetime of the rocket combustion chamber. This work presents the implementation of fully consistent real gas thermodynamics into the commercially available CFD-code ANSYS CFX. Special attention is paid to real gas mixing effects and their influence on CFD-simulations at transcritical conditions.

Modular divert and attitude control system

Abstract: There is disclosed a vehicle and methods for maneuvering the vehicle. The vehicle may include a plurality of multiple-impulse rocket motors, each of which comprises a plurality of independently ignitable solid fuel propellant charges, and a processor that generates at least one command to ignite at least one solid fuel propellant charge of at least one of the plurality of multiple-impulse rocket motors.

Optimization of a Hybrid Rocket Upper Stage with Electric Pump Feed System

Abstract: An electrical pump is considered to feed the oxidizer into the combustion chamber of a hybrid rocket motor, which is used as the third stage of a three-stage launcher. The motor uses hydrogen peroxide as the oxidizer and polyethylene as the fuel; advanced Lithium batteries are adopted to power the pump. The design of the hybrid rocket motor and the trajectory are simultaneously optimized by means of a nested direct/indirect procedure. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload for assigned characteristics of the first and second solid propellant stages and final orbit. The optimization provides the optimal values of the main engine design parameters, the corresponding grain and engine geometry, and the control law. A mission profile based on the Vega launcher is considered. The performance obtained using an electrical pump feed system is compared with pressuregas feed systems. Results show the relevant improvements that can be obtained when a present-technology electrical pump feed system is adopted.