Ducted Rocket Research Articles

Computing Method Investigation and Verification of Gas-Solid Combustion in Magnesium-Aluminum Based Propellant Ducted Rocket

The combustion mechanism, which consisting of 22 species and 23 reaction equations, and three discrete models such as inertia, combusting, modification droplet, are employed for the investigation of gas-solid combustion in magnesium-aluminum based propellant ducted rocket based on thermal performance calculation. And path lines, temperature distribution, sediments are discussed after the computing method is validated by direct-connect experimentation and the flow field information, which obtained by numerical method and coincided with currently conclusions. The results indicated that the proposed method is reliable and practicable.

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

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 buzz. First, the mathematical models involving gas flow regulating system and ducted rocket are introduced. The description of inlet buzz margin is also given and analyzed. Second, the multi-objective switching control problem of ducted rocket is proposed and discussed. Third, a robust bumpless transfer scheme for solving the problem is presented. The designed bumpless transfer compensator based on model reference adaptive sliding-mode control ensures that the switched system performs a smooth transition at the transfer moment. The stability analysis is then given utilizing the Lyapunov functional approach. Finally, the bumpless transfer strategy is applied to the ducted rocket control system, and the simulation results show its effectiveness.

Friction-Compensation Control of Gas-Flow Regulation for Ducted Rockets Based on Adaptive Dither Method

AbstractIn this paper, the friction generated in gas-flow regulation of ducted rockets is considered, and a compensation method is discussed to eliminate its effect. First, the dynamic mathematic model of gas flow–regulating system considering friction is established and analyzed. Then, the mechanism of friction-compensation method based on additional dither signal is proposed, and the compensation loop with adaptive-amplitude dither signal is added to the original system to overcome the effect of the varied friction. The simulation results show the validation of the adaptive dither signal in solving the problems caused by the variable friction in gas-regulation system of ducted rockets.

Boron Particle Ignition in Secondary Chamber of Ducted Rocket

In the secondary chamber of ducted rocket, there exists a relative speed between boron particles and air stream. Hence, the ignition laws under static conditions cannot be simply applied to represent the actual ignition process of boron particles, and it is required to study the effect of forced convective on the ignition of boron particles. Preheating of boron particles in gas generator makes it possible to utilize the velocity difference between gas and particles in secondary chamber for removal of the liquid oxide layer with the aid of Stoke's forces. An ignition model of boron particles is formulated for the oxide layer removal by considering that it results from a boundary layer stripping mechanism. The shearing action exerted by the high-speed flow causes a boundary layer to be formed in the surface of the liquid oxide layer, and the stripping away of this layer accounts for the accelerated ignition of boron particles. Compared with the King model, as the ignition model of boron particles is formulated for the oxide layer removal by considering that it results from a boundary layer stripping mechanism, the oxide layer thickness thins at all times during the particle ignition and lower the ignition time.

Thrust control system design of ducted rockets

The investigation of the thrust control system is aroused by the need for propulsion system of ducted rockets. Firstly the dynamic mathematical models of gas flow regulating system, pneumatic servo system and ducted rocket engine were established and analyzed. Then, to conquer the discussed problems of thrust control, the idea of information fusion was proposed to construct a new feedback variable. With this fused feedback variable, the thrust control system was designed. According to the simulation results, the introduction of the new fused feedback variable is valid in eliminating the contradiction between rapid response and stability for the thrust control system of ducted rockets.

Switching control of thrust regulation and inlet buzz protection for ducted rocket

The renewed interest on ducted rockets impulses their investigation. In this article, switching control in the working process of ducted rockets is focused on, in order to obtain optimal thrust control while avoiding phenomena like inlet buzz. Firstly multi-objective control problems of ducted rockets during its working process are discussed. Then the dynamic mathematical models of gas flow regulating system, thrust regulation control loop and inlet buzz protection loop are established and analyzed. Lastly, the switching strategy and PID controller are applied to the ducted rocket system, and the influence of integral limitation of controllers is analyzed. In conclusion, it is useful to introduce the multi-objective switching control to ducted rockets, and simulation results show its validity.

Supersonic inlet buzz margin control of ducted rockets

The maximum performance of a supersonic inlet will be achieved when operating as close as possible to its buzz boundary. In order to maintain high performance without crossing the buzz boundary, an active buzz margin predictor and controller is necessary. The goal of a control system is to acquire inlet buzz margin and maintain it as the designated value and predict inlet buzz before it occurs and then take some measures to buzz mitigation or inlet restart. The inlet buzz boundary and the margin of ducted rockets were first discussed and analysed. Then the dynamic mathematical model of a gas flow control system was established. Lastly, the inlet buzz margin controller of a ducted rocket was designed. Simulation results show that there exist anti-regulation characteristics for gas flow control for ducted rockets, and it leads to the particularity of inlet buzz margin control for ducted rockets in contrast to liquid ramjet. The designed inlet buzz margin has to be bigger to avoid the appearing of inlet buzz because of anti-regulation characteristics; otherwise, the inlet buzz phenomenon probably appears. If the inlet buzz of the ducted rocket appears, the command of decreasing fuel flow could not be given immediately; otherwise, the inlet buzz will aggravate.

Extended combustion model for single boron particles – Part I: Theory

Ramjet engines have significant advantages when compared to conventional rocket motors concerning specific impulse, manoeuvrability, and range. Boron particle addition to the propellant of ducted rockets further increases this potential due to a very high heating value. However, the combustion of boron particles is a very complex process because of an inhibiting oxide layer covering the particles. This layer has to be removed before vigorous combustion can start. The boron particle combustion process runs in two distinct stages. In the literature review presented in this article two combustion models for single boron particles are outstanding. A very detailed model by the Princeton/Aerodyne group features hundreds of elementary reactions and considers all physical processes in the particle environment. It is very elaborate and, thus, not suitable for incorporation into three-dimensional CFD-calculations at present. The second model developed at Penn State University takes on a global approach with only a few reactions which makes it promising for CFD applications. A careful analysis of this model revealed some inconsistencies, errors and drawbacks which gave rise to the new model presented in this paper. The new model comprises a consistent formulation of the heat and mass transfer processes in the particle environment based on a quasi-steady approach, accounts for boron evaporation which is a relevant process despite the high boiling point of boron, and it considers the influence of forced convection on the particle conversion. The chemical reaction rates adopted from the original model were revised and are slightly changed, the differential equations to be solved are corrected and an iterative solution algorithm is introduced. A careful validation of the model is presented in Part II of this paper showing that the new model is suitable for boron particle sizes relevant for ramjet combustion chambers.

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

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 for the ducted rocket was established, and comparison of the two regulating modes regarding non-linearity and uncertainty was done. In contrast with the area-regulated mode, the non-linearity of the pressure-balanced mode is smaller and the robustness is stronger. Then, the study of the the gas control system is performed based on the model established in this article. The introduction of high-gain feedback suppresses the effects of non-linearity and uncertainty in the pneumatic servo system. Also, a proportional-integral controller for the gas regulating system based on linear matrix inequality is designed. Lastly, an experimental investigation of the gas flow control system is done to verify the validity of the gas control system. The experimental result shows that the gas flow control system has good performance. The turndown ratio of mass flow through the control valve was up to 3.7:1.

Numerical analysis of heterogeneous combustion in a ducted rocket

A model to analyse the heterogeneous combustion process that takes place in an axisymmetric combustor of an air ducted rocket is presented, including radiation. The fuel, injected into the combustor through a simple injector, is a mixture of gaseous products, carbon particles, and MgO inert particles that results from the combustion of a solid fuel (HTPB-AP-Mg). The turbulence model used is the realizable k—π model. The adopted reaction model is a reduced kinetic scheme composed of five reactions, three in gaseous phase and two heterogeneous. The model ‘finite rate/eddy dissipation model of Magnussen and Hjertager’ is used to calculate the production terms in the gas phase, meanwhile the mixed model of Field, Baum, and Street is used for the oxidation reaction of carbon particles. After a detailed analysis of the numerical results, obtained with a commercial code for a specific combustor geometry and one set of boundary conditions both for the gaseous and solid phases, the existence of a special combustion pattern is found. It is composed of a turbulent diffusion jet flame attached to the injector, where only the gaseous fuel is burned, followed by a reaction region that envelops the jet of carbon particles at the aft of the combustion chamber. This jet of carbon particles oxidizes at a finite rate producing CO and CO2. The emission of CO and C(s) from the combustor is the main cause of high efficiency reduction. Radiation has little effect on calculated global variables, such as efficiency, mass flowrate, and total pressure loss, but on the contrary its effects on the temperature fields became important. Chilling of the gas phase was observed together with a reduction of solid particle temperature, however wall combustor temperature increased greatly. Simulation results are interpreted in the light of experimental observations and available numerical solutions in the literature.

ハイブリッドラムジェットの理論性能検討

Hybrid-ramjet is proposed in this paper. It is the combined engine of a liquid fuel ramjet and a ducted rocket. The high Isp performance and the wide fuel flow control range of a liquid fuel ramjet and the high combustion stability of a ducted rocket are expected. Theoretical calculations of the hybrid-ramjet which has same mass solid fuel and liquid fuel, were conducted. The results shows 50% higher Isp and twice fuel control range compared with a ducted rocket.

Performance Prediction of a Ducted Rocket Combustor Using a Simulated Solid Fuel

The ducted rocket is a supersonic flight propulsion system that takes the exhaust from a solid fuel gas generator, mixes it with air, and burns it to produce thrust. To develop such systems, the use of numerical models based on computational fluid dynamics (CFD) has been increasing, but to date only simplified treatments of the combustion within ducted rockets have been reported, likely due to the difficulties in characterizing and accurately modeling the partially reacted, particle-laden fuel exhaust from the gas generator. Through a careful examination of the governing equations and experimental measurements, a CFD-based methodology that properly accounts for the influence of the gas generator exhaust, particularly the solid phase, has now been developed to predict the performance of a ducted rocket combustor using a simulated solid fuel. It uses an equilibrium-chemistry probability density function combustion model with two separate streams, one gaseous and the other of 75-nm-diam carbon spheres, to represent the exhaust products from the gas generator. After extensive validation with direct-connect combustion experiments over a wide range of geometries and test conditions, this CFD-based method was able to predict, within a good degree of accuracy, the combustion efficiency of a ducted rocket combustor.

Chemical Kinetics and Heat Transfer Issues for a Safe Bench Scale Production of Partially Triazole Substituted Glycidyl Azide Polymer (GAP)

AbstractThe reaction of azido polymers with double and triple bonds of light olefins can lead to triazole and aziridine cycles formation on the azide bearing sites of the polymer. An experimental investigation of the burning behaviour of these modified polymers has shown that triazole substituted GAPs exhibit a higher burning rate at low pressure, which may be a desirable characteristic for ducted rocket applications. However, the chemical reaction of interest is highly exothermic and as such, it represents a significant safety risk when conducted in bulk. In order to reduce this risk, we have assessed the chemical kinetics of triazole formation by NMR spectroscopy and linked this information with DSC measurements to produce a safe procedure for the synthesis of GAP‐triazole polymer in a chemical reactor.

On the Optimization of GAP-based Ducted Rocket Fuels from Gas Generator Exhaust Characterization

The performance of a ducted rocket motor depends greatly on the composition of the exhaust products from its solid fuel gas generator. A series of experiments has been conducted to characterize the exhaust of GAP/carbon gas generator fuels as a function of pressure and grain formulation. Fuel samples were ignited in a closed vessel and gases collected for further analysis using gas chromatography. The major gaseous components identified in the exhaust stream were nitrogen, hydrogen and carbon monoxide while the condensed phase was found to be a carbon-rich soot. The relative amounts of each major component depended greatly on the chamber pressure and the precise formulation used, thus providing a simple means of optimizing the exhaust composition to maximize the performance of the ducted rocket propulsion systems.

FLOW VISUALIZATION OF FOUR-INLET DUCTED ROCKET ENGINE CONFIGURATIONS

Abstract : A water flow visualization facility was constructed in order to investigate the mixing processes inside of four-inlet ducted rocket engines (DRE) with various flow rates and geometries. The observed flow fields were documented with video and conventional photography so that relationships between various geometric/flow parameters and the flow field behavior could be obtained.

Energetic insensitive propellants for solid and ducted rockets

After stating that propellants used on operational missiles do not yet meet all insensitive munition (IM) recommendations of the official French document, Instruction Nr 2060 DGA, IPE, characteristics of new formulations of composites and low-signature propellants are discussed. Butacene®, a prepolymer with grafted combustion catalyst species, reduces the sensitivity and reactivity of high-burning rate composite propellants when submitted to thermal hazards (or threats) without any loss of energetic performance in the operational temperature range. For minimum-smoke propellants, the use of energetic binders places limitations on both the amount of explosive fillers (nitramine) and the sensitivity of the energetic plasticizers to avoid the bore effect (with delayed shock-to-detonation transition phenomena) and losses in performance may have to be allowed. Other results obtained on extruded double-base propellants (SD 1175 and SD 1178), and gas-generator propellants show that it may be possible to design a motor that meets all the MIL-STD 2105 IM requirements. It is concluded that some progress is expected to be made in the near future using new energetic molecules such as hexanitrohexaazaisowurtzitane (CL-20), polyazidoglycidvie and hydrazinium nitroformate (mainly in minimum smoke propellant) to produce both high performance and high immunity levels.

Computation of turbulent reacting flow in a solid-propellant ducted rocket

A mathematical model for computation of turbulent reacting flows is developed under general curvilinear coordinate systems. An adaptive, streamline grid system is generated to deal with the complex flow structures in a multiple-inlet solid-propellant ducted rocket (SDR) combustor. General tensor representations of the k-epsilon and algebraic stress (ASM) turbulence models are derived in terms of contravariant velocity components, and modification caused by the effects of compressible turbulence is also included in the modeling. The clipped Gaussian probability density function is incorporated in the combustion model to account for fluctuations of properties. Validation of the above modeling is first examined by studying mixing and reacting characteristics in a confined coaxial-jet problem. This is followed by study of nonreacting and reacting SDR combustor flows. The results show that Gibson and Launder`s ASM incorporated with Sarkar`s modification for compressible turbulence effects based on the general curvilinear coordinate systems yields the most satisfactory prediction for this complicated SDR flowfield. 36 refs.

Turbulent flows in two-dimensional and three-dimensional simulated SDR combustors

The mean velocity and turbulence quantities in simulated two-dimensional (2-D) and three-dimensional (3-D) solid-propellant ducted rocket (SDR) combustor flow fields were measured using a two-component laser-Doppler velocimetry system. The Reynolds number, based on the combustor height and bulk mean velocity and momentum ratio of the axial inlet to side inlets were 4.56 × 10 4 and 0.11, respectively, for the 2-D flow field and 5.92 × 10 4 and 0.20 for the 3-D flow field. It is found that the well-organized dome vortices and separating bubbles in the 2-D flow are absent in the 3-D flow. To provide a reference for turbulence modeling, the 2-D flow is also characterized in terms of turbulent structure parameter. In addition, the penetration of the axial jet through the side-inlet jets and the mixing of the axial-inlet with the side-inlet streams are documented.

Turbulent Flows in a Model SDR Combustor

An experimental study is reported on the isothermal flow fields in a model solid-propellant ducted rocket combustor with two opposing side inlets. The measurements were made by using a four-beam two-color laser-Doppler velocimeter (LDV). Three values of momentum ratio (Ma/Ms) of the axial- to side-inlet jet—0.025, 0.11, and 1.28—were selected to investigate their effects on the flow characteristics. The Reynolds number, based on the air density, combustor height, and bulk velocity, was 4.56 × 104. The flow field was characterized in terms of the mean-velocity vectors and contours, joint probability functions, mean reattachment lengths, spreading rate of the axial-jet width, and Reynolds stress and turbulence kinetic energy contours. The LDV measured mean reattachment lengths were found to well agree with the corresponding flow-visualization photograph. In addition, the three Ma/Ms values provided three characteristic flows which are useful in testing the computational models. Further, correlations between the present cold-flow and previous reading-flow studies were documented in detail. It was found that from the fluid dynamic point of view Ma/Ms = 0.11 was preferable to the other two values of Ma/Ms.

Turbulence/shock wave interactions in chemically reacting flows

Physical interactions between turbulence and shock waves are very complex phenomena. If these interactions take place in chemically reacting flows, the degree of complexity increases dramatically. Examples of applications may be cited in the area of supersonic combustion, in which the controlled generation of turbulence and/or large scale vortices in the mixing and flame-holding zones is crucial for efficient combustion. Equally important, shock waves interacting with turbulence and chemical reactions affect the combustor flowfield resulting in enhanced relaxation and chemical reaction rates. Chemical reactions in turn contribute to dispersion of shock waves and reduction of turbulent kinetic energies. Computational schemes to address these physical phenomena must be capable of resolving various length and time scales. These scales are widely disparate and the most optimum approach is found in explicit/implicit adjustable schemes for the Navier-Stokes solver. This is accomplished by means of the generalized Taylor-Galerkin (gtg) finite element formulations. Adaptive meshes are used in order to assure efficiency and accuracy of solutions. Various benchmark problems are presented for illustration of the theory and applications. Geometries of ducted rockets, supersonic diffusers, flame holders, and hypersonic inlets are included. Merits of proposed schemes are demonstrated through these example problems.