Showing papers on "Rocket published in 2013"

Numerical Analysis of Nozzle Material Thermochemical Erosion in Hybrid Rocket Engines

Abstract: Ablative materials are commonly used to protect the nozzle metallic housing and to provide the internal contour to expand the exhaust gases in both solid and hybrid rockets. Because of interaction with hot gas, these materials are chemically eroded during rocket firing, with a resulting nominal performance reduction. The objective of the present work is to study the erosion behavior of graphite nozzles in hybrid engines at different operating conditions and compare results with those obtained for solid motors. A main distinctive feature of hybrid engine operating conditions is, in fact, a greater concentration of oxygen-containing combustion products than solid motors. The adopted approach relies on a validated full Navier–Stokes flow solver coupled with a thermochemical ablation model that takes into account heterogeneous chemical reactions at the nozzle surface, rate of diffusion of the species through the boundary layer, ablation species injection in the boundary layer, heat conduction inside the nozzl...

Chemical erosion of carbon-phenolic rocket nozzles with finite-rate surface chemistry

Abstract: Ablative materials are commonly used to protect the nozzle metallic housing and to provide the internal contour to expand the exhaust gases in solid rocket motors. Because of the extremely harsh environment in which these materials operate, they are eroded during motor firing with a resulting nominal performance reduction. The objective of the present work is to study the thermochemical erosion behavior of carbon-phenolic material in solid rocket motor nozzles. The adopted approach relies on a validated full Navier–Stokes flow solver coupled with a thermochemical ablation model, which takes into account finite-rate heterogeneous chemical reactions at the nozzle surface, rate of diffusion of the species through the boundary layer, pyrolysis gas and char-oxidation product species injection in the boundary layer, heat conduction inside the nozzle material, and variable multispecies thermophysical properties. The results obtained with the proposed approach are compared with two sets of experimental data: subs...

Numerical Analysis of Hybrid Rocket Combustion

Abstract: Hybrid rocket combustion has important effects on rocket performance. The solid fuel regression rate is an important quantity in the hybrid rocket operation. In the past years, experimental and analytical investigations have been conducted to find correlations to correctly predict the regression rate. Numerical computations are becoming more important in the estimation of the characteristic parameters of such a complex combustion that embraces many different phenomena. This study predicts the regression rate of the Hydroxyl-Terminated Poly-Butadiene/Gaseous Oxygen formulation and its sensitivities to some operating parameters, such as combustion chamber pressure, oxygen inlet temperature, and mass flow rate. Furthermore, an analysis of other variables is used to explain the experimentally observed regression rate behavior. Particular emphasis is placed on the effect of the oxygen between the flame and the surface, which is considered responsible for the pyrolysis process enhancement.

Coupled Wall Heat Conduction and Coolant Flow Analysis for Liquid Rocket Engines

Abstract: Coolant flow modeling in regeneratively cooled rocket engines fed with turbo machinery is a challenging task because of the high wall temperature gradient, the high Reynolds number, the high aspect ratio of the channel cross section, and the heat transfer coupling with the hot-gas flow and the solid material. In this study the effect of wall heat conduction on the coolant flow is analyzed by means of coupled computations between a validated Reynolds-averaged Navier–Stokes equations solver for the coolant flowfield and a Fourier’s equation solver for the thermal conduction in the solid material. Computations of supercritical-hydrogen flow in a straight channel with and without coupling with the solid material are performed and compared to understand the role played by the coupling on the coolant flow evolution. Finally, the whole cooling circuit of the space shuttle main engine main combustion chamber is analyzed in detail and discussed for the sake of comparison of results obtained with the present couple...

Numerical and Experimental Investigation of Unidirectional Vortex Injection in Hybrid Rocket Engines

Abstract: A study on vortex injection in hybrid rocket engines with nitrous oxide and paraffin has been performed. The investigation followed two paths: first, the flowfield was simulated with a commercial computational fluid dynamics code; then, burn tests were performed on a laboratory-scale rocket. The computational fluid dynamics analysis had the dual purpose to help the design of the laboratory motor and to understand the physics underlying the vortex flow coupled with the combustion process compared with axial injection. Vortex injection produces a more diffuse flame in the combustion chamber and improves the mixing process of the reactants, both aspects concurring to increase the c* efficiency. A helical streamline develops downstream of the injection region, and the pitch is highly influenced by combustion, which straightens the flow due to the acceleration in the axial direction imposed by the temperature rise. Experimental tests with similar geometry have been performed. Measured performance shows an incr...

Electrical Solid Propellants: A Safe, Micro to Macro Propulsion Technology

Abstract: In this work we developed three high performance families of plastisol solid propellants that are IM-capable and highly manufacturable, using benign processes and “green” ingredients. Highly aluminized Electric Solid Propellant (ESP) “igniterless” rocket motors have been demonstrated at 4inch diameter scale. These igniterless rocket motors use inert metal electrodes for direct ignition of the electric solid propellant, thereby eliminating the need for pyrotechnic ignition systems. A fast cook-off Insensitive Munitions test of a 4-inch diameter aluminum-cased ESP motor yielded a TypeV reaction, with no materials ejected from the burn pit. We successfully demonstrated pyrogen-free, igniterless electrically-initiated ignition of 200 lbf thrust class static tests in three separate firings conducted during the development program. In related activities, both nonmetalized minimum signature formulations and reduced-signature boron-fueled plastisol ESP compositions continue to undergo qualification testing for technology transfer applications in defense and commercial endeavors.

Numerical analysis of fuel regression rate distribution characteristics in hybrid rocket motors with different fuel types

Abstract: This paper presents three-dimensional numerical simulations of the hybrid rocket motor with hydrogen peroxide (HP) and hydroxyl terminated polybutadiene (HTPB) propellant combination and investigates the fuel regression rate distribution characteristics of different fuel types. The numerical models are established to couple the Navier-Stokes equations with turbulence, chemical reactions, solid fuel pyrolysis and solid-gas interfacial boundary conditions. Simulation results including the temperature contours and fuel regression rate distributions are presented for the tube, star and wagon wheel grains. The results demonstrate that the changing trends of the regression rate along the axis are similar for all kinds of fuel types, which decrease sharply near the leading edges of the fuels and then gradually increase with increasing axial locations. The regression rates of the star and wagon wheel grains show apparent three-dimensional characteristics, and they are higher in the regions of fuel surfaces near the central core oxidizer flow. The average regression rates increase as the oxidizer mass fluxes rise for all of the fuel types. However, under same oxidizer mass flux, the average regression rates of the star and wagon wheel grains are much larger than that of the tube grain due to their lower hydraulic diameters.

Isocyanate‐Free and Dual Curing of Smokeless Composite Rocket Propellants

Abstract: Landsem, Eva; Jensen, Tomas Lunde; Kristensen, Tor Erik Holt; Hansen, Finn Knut; Benneche, Tore; Unneberg, Erik. Isocyanate-Free and Dual Curing of Smokeless Composite Rocket Propellants. Propellants, explosives, pyrotechnics 2013 ;Volum 38.(1) s. 75-86

Combustion Instability Mechanisms in a Pressure-coupled Gas-gas Coaxial Rocket Injector

Abstract: : An investigation of the instability mechanism present in a laboratory rocket combustor is performed using computational fluid dynamics (CFD) simulations. Three cases are considered which show different levels of instability experimentally. Computations reveal three main aspects to the instability mechanism, the timing of the pressure pulses, increased mixing due to the baroclinic torque, and the presence of unsteady tribrachial flame. The stable configuration shows that fuel is able to flow into the combustor continuously allowing continuous heat release. The unstable configuration shows that a disruption in the fuel flow into the combustor allows the heat release to move downstream and new fuel to accumulate in the combustor without immediately burning. Once the large amounts of fuel in the combustor burn there is rapid rise in pressure which coincides with the timing of the acoustic wave in the combustor. The two unstable cases show different levels of instability and different reignition mechanism.

A mechanism for unsteady separation in over-expanded nozzle flow

Abstract: Shock wave induced separation in an over-expanded planar nozzle is studied through numerical simulation. These Large-Eddy Simulations (LES) model previous experiments which have shown unsteady motion of the shock wave in flows with similar geometries but offered little insight into the underlying mechanism. Unsteady separation in nozzle flow leads to “side loads” in the rocket engine which can adversely affect the stability of the rocket. A mechanism for the low-frequency shock motion is identified and explained using the LES data. This mechanism is analyzed for a series of over-expanded planar nozzles of various area ratios and nozzle pressure ratios. The effect of grid resolution and Reynolds number on the instability is discussed. A simple reduced order model for the unsteady shock behavior is used to further validate the proposed mechanism. This model is derived from first principles and uses data from the LES calculations to capture the effects of the turbulent boundary layer and shear layer.

Performance of Dicyclopentadiene/H2O2-Based Hybrid Rocket Motors with Metal Hydride Additives

Abstract: An experimental performance evaluation of metal hydride fuel additives for hybrid rocket motor propulsion systems is examined in this paper. Some metal hydride additives offer improved performance, but a common issue is material aging. An accelerated aging study revealed that dicyclopentadiene can protect sodium borohydride (NaBH4) particles from exposure to air and water vapor much better than conventional hydroxyl-terminated polybutadiene. Static hybrid rocket motor experiments were conducted using dicyclopentadiene as the fuel. Sodium borohydride and aluminum hydride (AlH3) were examined as fuel additives. Ninety percent rocket-grade hydrogen peroxide was used as the oxidizer. In this paper, the sensitivity of solid fuel regression rate and characteristic velocity efficiency to total fuel grain port mass flux and particle loading is examined. Chamber pressure histories revealed steady motor operation in most tests, with reduced ignition delays when using the NaBH4 fuel additive. The addition of NaBH4 a...

Design of an Airbreathing Second Stage for a Rocket-Scramjet-Rocket Launch Vehicle

Abstract: The most promising alternative to rockets for improved access to space involves staged systems using airbreathing propulsion. With scramjet technology improving, a number of airbreathing assisted access-to-space vehicle concepts have recently been proposed, including a three-stage rocket-scramjet-rocket launch architecture for payload masses on the order of 100 kg. This article presents a design methodology developed for the airbreathing second stage of such a system. This methodology uses multidisciplinary design optimization with simplified methods for the calculation of vehicle aerodynamics, propulsion, and mass. It has been applied to the design of a reusable scramjet-powered winged cone vehicle with a near-term Mach 6–12 hydrogen-fueled scramjet for propulsion. Through the manipulation of five vehicle design parameters, including the size and position of the engines, and flying the vehicle along constant dynamic pressure trajectories, a configuration was developed to maximize payload mass fraction to...

3D fluid–structure interaction analysis of a typical liquid rocket engine cycle based on a novel viscoplastic damage model

Abstract: SUMMARY In many space missions, expandable or reusable launch systems are used. In this context, the reliable design of liquid rocket engines (LREs) is a key issue. In the present paper, we present a novel combination of numerical schemes. It is applied to model the extreme physical phenomena a typical LRE undergoes during its loading cycles. The numerical scheme includes a partitioned fluid–structure interaction (FSI) algorithm in combination with a unified viscoplastic damage model. This allows the complex description of the material response under cyclic thermomechanical loading taking place in LREs. In this regard, we focus on the response of the cooling channel wall that is made from copper alloys. For the coupled FSI analysis, the individual domains of the rocket thrust chamber are modeled by a 3D parametrized approach. The well-established single field solver codes, DLR TAU for the hot gas and ABAQUS FE software for the structural domain, are coupled via the inhouse developed simulation environment ifls. Ifls provides the necessary algorithms for a partitioned coupling approach such as individual code steering, data interpolation, time integration and iteration control. Finally, the results of an FSI analysis of a complete engine cycle are presented. They show the potential of the new numerical scheme for the lifetime prediction of LREs.Copyright © 2013 John Wiley & Sons, Ltd.

Numerical and Experimental Investigations of Epsilon Launch Vehicle Aerodynamics at Mach 1.5

Abstract: The Epsilon launch vehicle, successor of the M-V rocket, which conveyed “Hayabusa” is currently under development in Japan. The Epsilon is also designed for sending scientific satellites to outer space, and its first flight is scheduled to be in 2013. In this study, by conducting both numerical simulations and wind-tunnel tests, the aerodynamic characteristics and associated flow features of the Epsilon launch vehicle are extensively investigated at Mach 1.5. The results provided are axial/normal/side forces, pitching/yawing/rolling moments, detailed three-dimensional flow structure, along with effects of the Reynolds number (between wind-tunnel and flight conditions), skin stringers (small devices on the main body), and the difference from another configuration called “NextGenEpsilon”. This set of data includes unavailable ones at either the experiment standalone or the actual flight. Magnitudes of computed aerodynamic coefficients are in good agreement with the experiment and within the design criteria....

Real-Time Estimation of Time-Varying Bending Modes Using Fiber Bragg Grating Sensor Arrays

Abstract: An important challenge in launch vehicle simulation and control is created by the time-varying mass and inertia of the vehicle, as well as the consequent changes in modal frequencies and modal shapes of the structure as propellant is exhausted. Estimating modal information from a limited number of onboard sensors is inadequate for attitude control of a launch vehicle in real time, and the use of additional conventional sensors is unwarranted because of the mass penalty and complexity. This limitation has forced mission planners to base vehicle control schemes on pre-calculated modal information from finite element models. Recent advances in fiber Bragg grating (FBG) sensor technology enable locating a large number of sensing elements along a rocket's structure with a negligible mass penalty. This opens the path for real-time modal estimation and control. This paper presents a novel approach for the real-time estimation of mode shapes on a variable mass structure using FBG sensor arrays. The method is vali...

Demonstration of the improved rocket efficiency in direct-drive implosions using different ablator materials.

Abstract: The success of direct-drive implosions depends critically on the ability to create high ablation pressures (∼100 Mbar) and accelerating the imploding shell to ignition-relevant velocities (>3.7×10(7 ) cm/s) using direct laser illumination. This Letter reports on an experimental study of the conversion of absorbed laser energy into kinetic energy of the shell (rocket efficiency) where different ablators were used to vary the ratio of the atomic number to the atomic mass. The implosion velocity of Be shells is increased by 20% compared to C and CH shells in direct-drive implosions when a constant initial target mass is maintained. These measurements are consistent with the predicted increase in the rocket efficiency of 28% for Be and 5% for C compared to a CH ablator.

Improved Rocket Efficiency in Direct-Drive Implosions Using Different Ablator Materials

Abstract: The success of direct-drive implosions depends critically on the ability to create high ablation pressures (∼100 Mbar) and accelerating the imploding shell to ignition-relevant velocities (>3.7×10(7 ) cm/s) using direct laser illumination. This Letter reports on an experimental study of the conversion of absorbed laser energy into kinetic energy of the shell (rocket efficiency) where different ablators were used to vary the ratio of the atomic number to the atomic mass. The implosion velocity of Be shells is increased by 20% compared to C and CH shells in direct-drive implosions when a constant initial target mass is maintained. These measurements are consistent with the predicted increase in the rocket efficiency of 28% for Be and 5% for C compared to a CH ablator.

Investigations on the Turbulent Wake of a Generic Space Launcher Geometry in the Hypersonic Flow Regime

Abstract: The turbulent wake flow of generic rocket configurations is investigated experimentally and numerically at a freestream Mach number of 6.0 and a unit Reynolds number of 10 · 10 6 .T heflow condition is based on the trajectory of Ariane V at an altitude of 50km, which is used as baseline to address the overarchingtasks of wakeflows in the hypersonicregimelikefluid-structuralcoupling,reversehot jets and base heating. Experiments using pressure transducers and high-speed schlieren measurement technique were conducted to gain insight into the local pressure fluctuations on the base and the oscillations of the recompression shock. This experimental configuration features a wedge-profiled strut orthogonally mounted to the main body. Additionally, the influence of cylindrical nozzle extensions attached to the base of the rocket is investigated, which is the link to the numerical investigations. Here, the axisymmetric model possesses a cylindrical sting support of the same diameter as the nozzle extensions. The sting supportallows investigationsof a undisturbedwakeflow. A time-accurate zonal RANS/LES approachwas applied to identify shocks, expansion waves, and the highly unsteady recompression region numerically. Subsequently,experimentaland numericalresults in the strut-avertedregionare opposed with regardto the wall pressure and recompression shock frequency spectra. For the compared configurations, experimental pressure spectra exhibit dominant Strouhal numbers at about SrD = 0.03 and 0.27 and the recompression shock oscillates at 0.2. In general, the numerical pressure and recompression shock fluctuations agree satisfactorily to the experimental results. The experiments with a blunt base reveal base-pressure spectra with dominant Strouhal numbers at 0.08 at the center position and 0.145, 0.21 − 0.22 and 0.31 − 0.33 at the outskirts of the base.

Multiphase Modeling of Water Injection on Flame Deflector

Abstract: This paper describes the use of an Eulerian Dispersed Phase (EDP) model to simulate the water injected from the flame deflector and its interaction with supersonic rocket exhaust from a proposed Space Launch System (SLS) vehicle. The Eulerian formulation, as part of the multi-phase framework, is described. The simulations show that water cooling is only effective over the region under the liquid engines. Likewise, the water injection provides only minor effects over the surface area under the solid engines.

Multi-objective regulating and protecting control for ducted rocket using a bumpless transfer scheme

Abstract: This article deals with the problem of multi-objective regulating and protecting control for a ducted rocket, in order to get maximum thrust while avoiding extremely dangerous phenomenon like inlet...

Flow Separation in Rocket Nozzles – An Overview

Abstract: The paper includes a literature overview on the history of separated rocket nozzle flow research as well as a data base on flow separation in convergent-divergent nozzles