Showing papers on "Rocket published in 2006"

Influence of a Conical Axial Injector on Hybrid Rocket Performance

Abstract: This paper was aimed at analyzing the static engine firings results obtained by means of a hybrid rocket where gaseous oxygen was supplied into axial-symmetric polyethylene cylindrical grains through two different injector configurations: an axial conical subsonic nozzle and a radial injector. The axial injector is considered interesting because of its easy design and the remarkable feature that it produces no significant pressure oscillations for the stabilizing effect due to the hot gas recirculation zone established within the combustion port. To take advantage of its qualities, the assessment of the regression rate variations under the flow field generated by this configuration is required. For the investigated set of operating conditions, the instantaneous regression rates exhibit a time-and-space dependence caused by the impinging jet zone dynamics, while the average regression rates are higher and less mass flux dependent than those achieved with the radial injection motor and expected from the turbulent flow through pipes. A comparison to the data from the radial injector was further drawn in terms of combustion efficiency, fuel consumption profiles, and combustion stability. The radial injector, at the same mass flux and pressure, produces lower regression rates, high pressure oscillations and worse combustion efficiency at the same L*, but more uniform fuel consumption.

Nanoparticles for solid rocket propulsion

Abstract: The characterization of several differently sized aluminium powders, by BET (specific surface), EM (electron microscopy), XRD (x-ray diffraction), and XPS (x-ray photoelectron spectroscopy), was performed in order to evaluate their application in solid rocket propellant compositions. These aluminium powders were used in manufacturing several laboratory composite solid rocket propellants, based on ammonium perchlorate (AP) as oxidizer and hydroxil-terminated polybutadiene (HTPB) as binder. The reference formulation was an AP/HTPB/Al composition with 68/17/15% mass fractions respectively. The ballistic characterization of the propellants, in terms of steady burning rates, shows better performance for propellant compositions employing nano-aluminium when compared to micro-aluminium. Results obtained in the pressure range 1-70 bar show that by increasing the nano-Al mass fraction or decreasing the nano-Al size, larger steady burning rates are measured with essentially the same pressure sensitivity.

An Active Diagnostic System for Structural Health Monitoring of Rocket Engines

Abstract: An active diagnostic system using built-in piezoelectric actuator/sensor networks was developed for monitoring crack growth in a rocket engine pipe. The diagnostic system combines a sensor network, portable diagnostic hardware and data analysis software which allows for real-time in situ monitoring and long term tracking of the structural integrity of pressure vessels. Experimental data shows that the system can detect a surface crack as small as 4 mm and a through-crack as small as 2 mm in the high pressure engine pipe made of Inconel 718. It was found that the actuator-sensor paths that are most sensitive to crack growth are the ones in which the crack is growing away from, rather than towards the path. This discovery will provide important guidelines for the design of a sensor network for crack detection. It was also observed that the bending mode (equivalent to the A0 mode in plates) is more sensitive than the longitudinal mode (equivalent to the S0 mode in plates) to crack growth.

Numerical studies on specific impulse of partially filled pulse detonation rocket engines

Abstract: Performance analyses of pulse detonation rocket engines (PDREs) were numerically studied, focusing on partialfill effects at ground tests. The initial detonable mixture, inert gas, fuel-fill fraction, equivalence ratio, and initial temperature of inert gas were changed as governing parameters. The simulation results were compared against those of previous studies and agreed well with them. The simulation results indicated that the initial mass fraction of the detonable mixture to total mass of the gas was the predominant factor for the specific impulse of partially filled PDREs. Based on the numerical results, a new, simple empirical formula is proposed to predict the specific impulse of partially filled PDREs.

Simulations of Solid-Propellant Rockets: Effects of Aluminum Droplet Size Distribution

Abstract: Flow modeling and simulation of solid-propellant rockets from first principles is quite challenging with several physical problems, including complex evolving geometries, turbulence, and multiphase flow with a chemically reactive disperse phase. To this end, a flexible simulation framework has been developed, in which multiphase flow computations are performed that include three-way coupling between phases (mixture-droplet-smoke), conservative coupling approach, and full heat release for the burning mechanisms. Results obtained from computations with burning aluminum droplets generating aluminum-oxide smoke are described for a generic rocket geometry. The effects of injected droplet size distribution obtained with two models are investigated and show the sensitivity of these distributions to the chamber flow dynamics, primarily at the nozzle inlet. The residence time and burning droplet diameter are verified by comparison with simple analytical predictions.

Six -DOF Modeling and Simulation of a Generic Hypersonic Vehicle for Control and Navigation Purposes

Abstract: This paper covers the development of a Six Degrees of Freedom simulation of a generic hypersonic vehicle (GHV) [1] based on two different aerodynamic models and the aerodynamic database developed in reference [2] for control and navigation purposes. The results from APAS, which is an engineering level CFD program, from a high fidelity CFD code, STARS, and from wind tunnel experiments are used in this research. For the GHV model a combined cycle engine including a turbojet, ramjet-scramjet, and rocket engines are designed to cover subsonic, supersonic, and hypersonic speeds. Using numerical linearization techniques including the Jacobian method, the LTI state equations were developed. This work is supported partially by an Air Force grant, in support of their emphasis on new aerospace vehicle concepts and hypersonic technologies.

Fuel Regression Rate Characterization Using a Laboratory Scale Nitrous Oxide Hybrid Propulsion System

Abstract: The purpose of this study was to characterize the regression rates of three traditional hybrid rocket fuels along with one novel fuel using nitrous oxide as the oxidizer. In order to complete these tests, a robust test facility was developed for use in this study as well as in future hybrid rocket programs at Stanford University. This stand allowed for rapid successive hot fire testing and the capability to test multiple fuel grain configurations. To date it has been used to conduct 40 hot fire tests. Performance values and fuel regression rate data are presented for HTPB, PMMA, HDPE, and sorbitol.

Observations of Triboelectric Charging Effects on Langmuir-Type Probes in Dusty Plasma

Abstract: [1] Investigation of Earth's mesosphere using sounding rockets equipped with a myriad of instruments has been a highly active field in the last 2 decades. This paper presents data from three separate instruments: an RF impedance probe, a DC fixed bias Langmuir probe, and an electric field probe, that were flown on a mesospheric sounding rocket flight investigating the presence of charged dust within and/or around a sporadic metal layer. The combined data set indicates a case of payload surface charging, the causes of which are investigated within this paper. A generic circuit model is developed to analyze payload charging and behavior of Langmuir-type instruments. The application of this model to the rocket payload indicates that the anomalous charging event was an outcome of triboelectrification of the payload surface from neutral dust particles present in the Earth's mesosphere. These results suggest caution in interpreting observations from the Langmuir class of instrumentation within dusty environments.

Single-stage hypersonic vehicle featuring advanced swirl combustion

Abstract: A single-stage hypersonic vehicle is comprised of a low-speed and a high-speed propulsion system. The low-speed propulsion system propels the single-stage vehicle to a threshold velocity, after which the high-speed propulsion system then takes over. The low-speed propulsion system includes a combined-cycle engine featuring a swirl generator that is integrated into a turbojet engine to provide a compact turbojet and swirl afterburner-ramjet propulsion system. The high-speed propulsion system includes a hypersonic engine that is operable at the threshold takeover velocity and beyond. In various embodiments, the high-speed propulsion system comprises a scramjet, rocket, or scramjet/rocket engine depending requirements. Benefits of the swirl generator design include its ability to rapidly and efficiently atomize, vaporize, mix and burn the fuel and oxidizer for the low speed propulsion system, significantly reduce length, weight, cooling requirements and complexity for both propulsion systems, while maintaining high propulsion performance and reducing propulsion and launch costs.

Fluid-Throat-Induced Shock Waves During the Ignition Transient of Solid Rockets

Abstract: PREDICTION and control of pressure and pressure-rise rate during the ignition transient of solid-propellant rocket motors with a nonuniform port are of topical interest. In certain designs, an ignition pressure spike and a high rate of pressure rise may adversely affect the steadiness and stability of burning, thermoviscoelastic response of the grain and inhibitors, and the dynamic response of the hardware parts.1 An excessive pressurization rate can cause a failure even when the pressure is below the design limit.2,3 Although, a great deal of research has been done in the area of solid rocket motors (SRMs) for more than six decades, the accurate prediction of the ignition transient in ports of high-performance solid rocket, with sudden expansion and/or steep divergence/convergence or protrusions has not previously been accomplished.

The Shefex Flight Experiment - Pathfinder Experiment for a Sky Based Test Facility

Abstract: On Thursday, October 27th, 2005 the SHarp Edge Flight EXperiment SHEFEX has been successfully launched at the Andoya Rocket Range in northern Norway. The project, being performed under responsibility of the German Aerospace Center (DLR) flew on top of a two-stage solid propellant sounding rocket. One purpose of the experiment is the investigation of possible new shapes for future launcher or re-entry vehicles applying a shape with facetted surfaces and sharp edges and to enable the time accurate investi-gation of the flow effects and their structural answer during the hypersonic flight from 90 km down to an altitude of 20 km. Additionally, the SHEFEX project is a starting point for a series of experiments which enable the acquisition of important knowledge in hypersonic free flight experimentation and which are an excellent test bed for new technological concepts. The present paper gives an overview about the philosophy and the layout of ex- periment and introduces preliminary outcomes of the post-flight analysis.

Steady Flows in the Slender, Noncircular, Combustion Chambers of Solid Propellant Rockets

Abstract: An analysis of the steady flows with axial vortices in slender, cylindrical, nonaxisymmetric cavities generated by injection from their porous walls is presented. This problem is encountered in the description of the flow in slender combustion chambers of solid-propellant rocket motors associated with the gasification of the solid propellant surrounding the combustion chamber. Nonreacting flow can be described in terms of self-similar solutions of the Navier-Stokes equations and the solution, calculated numerically for noncircular grain configurations, shows strong axial vortices, with a viscous vortex core that has been analyzed asymptotically for large Reynolds numbers. It was found that the important property of the flow, namely, C 1/2 S/Π, where C is a constant determining the axial pressure gradient, S and n are the cross section area and its perimeter, respectively, becomes unexpectedly close to π/2 at large Reynolds numbers, independently on the geometry of the cross section of the cavity. One can suppose that C 1/2 S/Π = π/2 for any cavities in the invicsid limit, a conjecture that also obtained a support from numerical calculations of flows in rectangular cavities generated by injection from their porous walls.

On-board Strain Measurement of a Cryogenic Composite Tank Mounted on a Reusable Rocket using FBG Sensors:

Abstract: This article presents the real-time strain measurement of a composite liquid hydrogen (LH2) tank using fiber Bragg grating (FBG) sensors. The tank was composed of carbon fiber reinforced plastic (C...

Research on Theory and Application of Multidisciplinary Design Optimization of Flight Vehicles

Abstract: With the increasing of the complexity of the flight vehicles, the Multidisciplinary Design Optimization (MDO) gets more and more attentions. In 1998, the Multidisciplinary Aerospace Design Optimization Research Center (MADORC) of National University of Defense Technology (NUDT) was built. In this paper, the researches done in MADORC are systematically introduced. Firstly, the definition of the MDO is discussed. The definition given here can be simply summed as: MDO=MD+MA+MO. Secondly, state-of-the-art of MDO of Flight vehicles is introduced and is classified as three generations. The characteristics of the three generations are compared. Thirdly, the research on MDO theory is present. They include the MDO-oriented Modeling, Design-Oriented Analysis, Approximation Method, Sensitivity Analysis, Design Space Search, Optimization Procedure, Information Infrastructure and MDO Frame. Fourthly, the research on MDO application is present, they include: (1) optimization of the interception missile; (2) integrated design of a ballistic missile; (3) integrated optimization of an air-to-air missile(AAM) and the solid propellant rocket ramjet; (4) MDO of Reusable Launch Vehicle (RLV); (5) MDO of satellite system. They belong to 3 generation of MDO application respectively.

Remotely controlled vehicle

Abstract: A remotely-controlled vehicle is disclosed. According to various embodiments, the vehicle includes components that may be added to an existing vehicle or added at the time of manufacture. The components include at least one of a pneumatic projectile launcher and a water cannon, a rocket launcher having one or more rockets launchable from the vehicle, at least one video camera system for capturing and transmitting video images, and a controller for controlling the projectile launcher, water cannon, rocket launcher, and video camera system. Each of the one or more rockets may include at least one solid-propellant rocket motor, and each of the one or more video camera systems may include at least one video camera mounted for selective orientation in at least one plane.

The SpaceLiner Concept and its Aerothermodynamic Challenges

Abstract: This paper outlines the technical concept of a high-speed intercontinental passenger transport using rocket based, suborbital launchers The paper gives an overview on the recent conceptual design status of the DLR SpaceLiner presenting geometrical size and mass data and describing results of trajectory simulations The rockets are based on an advanced but technically conservative approach not relying on exotic technologies The two-stage, fully reusable vehicle is designed as an “exceedingly reliable” system to overcome the safety deficits of current state-of-the-art launchers An otherwise advantageous skipping trajectory is generating very high heat loads in critical areas which could be successfully addressed by advanced ceramic materials with active cooling DLR is preparing a fundamental research test campaign on active leading edge transpiration cooling in high enthalpy flow DLR's arc heated, high enthalpy facility foreseen for the experiments is described in the paper The advanced C/C-structures, the selection of the cooling fluid, and the design of the experimental models are briefly discussed Supported by the results of the technical study, some programmatic issues concerning space flight are discussed A preliminary cost analysis of the rocket powered vehicle gives an indication on the economic feasibility and its impact on launch vehicle production The development of very high speed intercontinental passenger transport might enable, as a byproduct, to considerably reduce the cost of space transportation to orbit

Examination of Mode Shapes in an Unstable Model Rocket Combustor

Abstract: ** † ‡ Experimental data from an unstable model rocket combustor are analyzed using a linear one-dimensional acoustic model. The focus is on determining the influence of the oxidizer tube of the swirl-coaxial injector element on the acoustics and stability of the system at varying chamber lengths. The most significant effects of the oxidizer tube are seen in the system resonance frequencies and amplitude of acoustic pressure oscillations in the combustion chamber. Resonance frequencies are shown to be non-integer multiples, unlike traditional longitudinal modes, at certain combustion chamber lengths and are verified by comparisons with experimental data. In addition, certain modes calculated by the acoustic model to be highly damped in the combustion chamber are found to be unstable during the experiment, thereby indicating that the unsteady combustion process is dominating the stability response. Further, comparisons of the acoustic model with detailed CFD analysis verify the mode shape predictions and indicate the additional influence of mean flow Mach number and nonlinear effects on the mode shapes and frequencies.

Rotational flow in tapered slab rocket motors

Abstract: Internal flow modeling is a requisite for obtaining critical parameters in the design and fabrication of modern solid rocket motors. In this work, the analytical formulation of internal flows particular to motors with tapered sidewalls is pursued. The analysis employs the vorticity-streamfunction approach to treat this problem assuming steady, incompressible, inviscid, and nonreactive flow conditions. The resulting solution is rotational following the analyses presented by Culick for a cylindrical motor. In an extension to Culick’s work, Clayton has recently managed to incorporate the effect of tapered walls. Here, an approach similar to that of Clayton is applied to a slab motor in which the chamber is modeled as a rectangular channel with tapered sidewalls. The solutions are shown to be reducible, at leading order, to Taylor’s inviscid profile in a porous channel. The analysis also captures the generation of vorticity at the surface of the propellant and its transport along the streamlines. It is from t...

Power for a space plane.

Abstract: This article discusses the development of an aircraft which can leave and enter the Earth's atmosphere without the aid of rocket propulsion "Scramjets" are designed to scoop oxygen from the atmosphere as they ascend. The weight savings from capturing air in flight instead of carrying it means that these jets can deliver approximately four times the thrust of a rocket for every pound of propellant consumed.

Investigation Of Sub- And Supercritical LOX/Methane Injection Using Optical Diagnostics

Abstract: The DLR Institute of Space Propulsion at Lampoldshausen has been working on dierent aspects of oxygen/methane combustion for a couple of years. Within this framework, the European High Pressure Research and Technology Test Facility P8 was equipped with a methane fluid system as a first step to enable single injector investigations of high pressure methane injection and combustion processes. The fluid system provides ambient temperature methane at flow rates up to 1 kg/s and at combustion chamber pressures up to 10 MPa. The measurement system has been adapted to the new fluid system to provide a precise determination of injection conditions. Hence, the P8 now allows investigations of LOX/CH4 combustion chamber processes under realistic rocket engine conditions. For the first operational tests, a single injector windowed rocket combustion chamber was operated at various steady-state conditions in the sub-, near- and supercritical regime with respect to the critical pressure of oxygen. The propellants were injected through a single shear coaxial injector element at temperatures of about 120 K and 275 K, respectively. High speed optical diagnostic techniques such as flame emission spectroscopy and shadowgraphy have been applied in the near injector region to assess injection and atomization behavior as well as flame anchoring and stabilization. The paper describes the modifications of fluid and MCC systems at P8 and presents preliminary results of the first test campaign.

Trajectories for Human Missions to Mars, Part 2: Low-Thrust Transfers

Abstract: We compute optimal low-thrust transfers (with constant thrust and constant specific impulse) between Earth and Mars over a range of flight times (from 120 to 270 days) and launch years (between 2009 and 2022). Unlike impulsive transfers, the mass-optimal trajectory depends strongly on the thrust and specific impulse of the propulsion system. A low-thrust version of the rocket equation is provided where the initial mass or thrust may be minimized by varying the initial acceleration and specific impulse for a given power-system specific mass and for a trajectory time of flight. With fixed time-of-flight transfers there is a minimum thrust and a maximum allowable specific mass; that is, if the available thrust is too low or the specific mass is too large then the desired transfer does not exist. We find the minimum allowable thrust for constrained time-of-flight missions is on the order of a few Newtons per metric ton of payload for power systems with tens of kg/kW. As expected, the thrust and AV requirements of the trajectories decrease with increasing flight times. By extending the flight time from 180 to 270 days the A V is reduced by 40% for powered captures and up to 35% for aeroassisted capture trajectories.

Characterisation of current pulses superimposed on the ontinuous current in upward lightning initiated from tall objects and in rocket triggered lightning

Abstract: Characterisation of current pulses superimposed on the ontinuous current in upward lightning initiated from tall objects and in rocket triggered lightning

Low-frequency electrostatic waves in the ionospheric E-region: a comparison of rocket observations and numerical simulations

Abstract: . Low frequency electrostatic waves in the lower parts of the ionosphere are studied by a comparison of observations by instrumented rockets and of results from numerical simulations. Particular attention is given to the spectral properties of the waves. On the basis of a good agreement between the observations and the simulations, it can be argued that the most important nonlinear dynamics can be accounted for in a 2-D numerical model, referring to a plane perpendicular to a locally homogeneous magnetic field. It does not seem necessary to take into account turbulent fluctuations or motions in the neutral gas component. The numerical simulations explain the observed strongly intermittent nature of the fluctuations: secondary instabilities develop on the large scale gradients of the largest amplitude waves, and the small scale dynamics is strongly influenced by these secondary instabilities. We compare potential variations obtained at a single position in the numerical simulations with two point potential-difference signals, where the latter is the adequate representation for the data obtained by instrumented rockets. We can demonstrate a significant reduction in the amount of information concerning the plasma turbulence when the latter signal is used for analysis. In particular we show that the bicoherence estimate is strongly affected. The conclusions have implications for studies of low frequency ionospheric fluctuations in the E and F regions by instrumented rockets, and also for other methods relying on difference measurements, using two probes with large separation. The analysis also resolves a long standing controversy concerning the supersonic phase velocities of these cross-field instabilities being observed in laboratory experiments.

SHEFEX - Hypersonic Re-entry Flight Experiment Vehicle and Subsystem Design, Flight Performance and Prospects

Abstract: The purpose of the Sharp Edge Flight Experiment (SHEFEX) was to investigate the aerodynamic behavior and thermal problems of an unconventional shape for re-entry vehicles comprising multi-facetted surfaces with sharp edges. The main object of this experiment was the correlation of numerical analysis with real flight data with respect to the aerodynamic effects and structural concept for the thermal protection system (TPS). The Mobile Rocket Base of the German Aerospace Center (DLR) was responsible for the test flight of SHEFEX on a two-stage unguided solid propellant sounding rocket, which was required to provide a velocity of the order of Mach 7. The SHEFEX vehicle was launched on the 27th of October 2005 from the Andoya Rocket Range, Norway. This paper presents the main design features of the vehicle and subsystems, the flight performance and the current plans for our next hypersonic project.

Rocket-propelled grenade protection system

Abstract: It comprises of a protective grid (12) of special configuration and hardness. In the interspaces (19) of the grid (12) there exists a polyester foam (13) for its adequate electrical insulation. The grid (12) possesses a serrate conductive entrance (14) on the side of rocket impact (40) and a conductive cathode (15) on its other side. All the electrodes (23) are connected with an electronic control unit (16) in order for the data to be transferred from the impact surfaces to this unit. The unit (16) is programmed to terminate the electric energy produced by the piezoelectric crystal to an ignition head of the Rocket Propelled Grenade RPG when this rocket strikes against the conductive entrance (14) of the protective grid (12). It can be used for the protection against the lethal Rocket propelled grenades RPG.