Ducted Rocket Research Articles

Dynamic Testing of a Hybrid-Propellant Water-Breathing Ram Rocket in Underwater Cruise Conditions

High-speed submerged marine vehicles, such as torpedoes, traveling at velocities of an order of 100 m/s and above, require powerful propulsion to overcome the tremendous hydrodynamic drag. This paper aims to investigate a marine hybrid-propellant water-breathing ram rocket (marine ramjet or ducted rocket) under various underwater cruise conditions. At high underwater cruise speeds, the ram rocket outperforms regular rocket motors, substantially increasing its specific impulse and thrust. This investigation utilized a unique test facility capable of dynamically testing the marine ramjet. Over 20 dynamic experiments have been conducted, revealing the submerged motor characteristics at different cruise speeds, water-to-propellant mass ratios, and oxidizer-to-fuel mass ratios, thereby creating a performance map of the marine ramjet. The results were compared with static firing data and theoretical calculations, showing a good agreement with standard specific impulse improvement of about 55% compared to a regular hybrid rocket, reaching a maximum value of 380 s. The significant increase in performance demonstrates the potential of the water-breathing ramjet for propelling high-speed underwater vehicles.

Stochastic Equilibrium Analysis to Optimize Solid Propellant Combustion Gas Composition

Solid fuel ducted rockets operate with two-stage combustion. The first stage generates burnt products from the combustion of a fuel-rich solid propellant. In the second stage, those combustion products react with air and are ejected through a nozzle to produce thrust. The combustion efficiency of such a device strongly influences its performance and depends on the composition of the rich-fuel-burnt pyrolysis products. This study investigates the sensitivity of carbon, hydrogen, oxygen, and nitrogen (CHON) solid fuels’ burnt product composition to the fuel properties, composition, and heat of formation. The calculations are performed assuming adiabatic, isobaric chemical equilibrium on synthetic species. Existing species’ properties constrain the properties of the synthetic species. The trends of formation of the main species observed in gas generators, i.e., H2, CH4, CO, N2, CO2, H2O, and solid carbon, are presented. The sensitivity of the molar concentration ratio ξ of H2 to CH4 to the oxygen balance and propulsive properties is compared with an example of gas generator propellants. Recommendations for the possible optimization of fuel composition are formulated. For blends of the existing CHON species reported in the literature, it is shown that improving ξ is only possible by degrading the specific impulse.

Evaluation of Boron Combustion for Ducted Rocket Applications Using Condensed Product Analysis

Boron, a metalloid, produces high energy upon combustion. It is recommended as an ingredient for fuel or propellant in rocket propulsion, despite the challenge of extracting its full thermal energy. So far no one has claimed the complete energy conversion of boron upon combustion. On the other hand, the current propulsion system of the Meteor missile uses boron-loaded propellant. The boron-loaded propellant provides an approximately three-fold increase in specific impulse compared to conventional propellants. The present study focuses on boron-HTPB-based solid fuels impregnated with early ignited particles as additives, aiming to assess the combustion performance of boron particles. These additives are magnesium (Mg), titanium (Ti), and activated charcoal (C), and their effects are evaluated based on the residual active boron content in the condensed combustion products (burned residues). An economical tool commonly called stagnation flow or opposed flow burner (OFB) is used to deflagrate the fuel sample by means of pressurized oxygen gas. The condensed combustion products are examined using a field emission scanning electron microscope (FESEM), X-ray diffraction (XRD), thermogravimetric analysis (TGA), and differential thermal analysis (DTA). Among the fuel combinations investigated here, magnesium has been found to be a good burning enhancer of boron, leaving the lowest active boron content (30%) in the burned-out residue. The current research aims to develop an efficient boron-containing solid fuel for hybrid propellant ducted rocket engine applications.

Effect of synergistic impact gas injection device on secondary combustion in boron-based ducted rockets

The combustion efficiency of the afterburner has an important effect on the overall performance of the ducted rocket. In some difficult working conditions with low incoming air temperature or high flight altitude, it is more difficult to achieve efficient combustion in the afterburner. Therefore, this paper mainly solves the problem of low primary gas combustion efficiency in the afterburner of boron-based ducted rockets when the incoming air temperature is low. The flow field can be improved by optimizing the gas injection mode, so as to improve the combustion efficiency. In this paper, a synergistic impact gas injection device associated with angles of two directions is designed. By controlling the self-impact angle α of gas/gas, and the mutual-impact angle β of gas/air, efficient mixing and combustion is realized. Through numerical simulation, the mixing degree and combustion efficiency, as well as the contours of B2O3 mass fraction are obtained, and the influence regularities and flow field characteristics are analyzed. The selection scheme of the optimal injection device is verified by a ground direct-connected experiment. Finally, through the jet observation experiments, the gas ejection form from the synergistic impact gas injection device is observed. The research of this paper shows that when the mutual-impact angle β is small, the larger the self-impact angle α is, the higher the mixing degree and combustion efficiency will be. When the mutual-impact angle β is large, each self-impact angle α has a high mixing degree and combustion efficiency. When the self-impact angle α is 40° or the mutual-impact angle β is 35°, the mixing can be completed in advance before the gas reaches the afterburner outlet. The higher the average mixing degree, the higher the combustion efficiency. This paper provides an innovative and efficient method for promoting combustion in the afterburner.

A short review of challenges and prospects of boron-laden solid fuels for ramjet applications

This short review article explores the advancement, challenges, and prospects of boron-laden solid fuel for ramjet. The energetic aspect of boron combustion is well known, but the complications associated with its efficient combustion persist and must be overcome to harness its potential completely. The strategies like coating, adding additives, and utilizing boron nanoparticles, which may improve its combustion performance, are widely studied. The latest development in advanced rocket propulsion systems, i.e., using a hybrid fuel-based gas generator for ducted rockets or air augmentation of hybrid rockets, seems a feasible idea; as it can provide two-stage combustion, which may significantly promote and improve the combustion performance of boron hence the specific impulse. Since it utilizes hybrid fuel, it retains its advantages of it as well as its drawbacks. To mitigate these drawbacks, various strategies/suggestions are briefly discussed here. A fundamental aspect of boron combustion must be fully understood and addressed for its applications in advanced propulsion systems.

Effect of gas injection holes on mixing and combustion in solid propellant ducted rockets

Combustion efficiency is an important parameter of the combustion chamber. Gas injection holes have an important influence on the mixing and combustion in solid propellant ducted rockets, but there is currently little research on gas injection holes. This paper describes numerical simulations of solid propellant ducted rocket combustion chambers, with the numerical method verified through a direct-connected experiment. We investigate the effect of adding injection holes to the gas generator nozzle on the mixing and combustion performance, and examine the mixing degree, combustion efficiency, and flow field characteristics. This paper focuses on the number and the position of the gas injection holes. The results show that, when the total mass flow rate is constant, the mixing effect becomes better as the dispersion degree of gas injection increases and the mass flow rate of each hole decreases. For this study, the 5-hole scheme has the best performance, and when the total mass flow rate and the number of injection holes are fixed, a smaller mass flow rate of gas injected toward the outlet of each air inlet is more conducive to better mixing and combustion effects. When the total mass flow rate is constant and the shape of the injection holes are complete circles, increasing the number of holes is more effective than decreasing the mass flow rate of gas injected toward the outlet of the air inlets in improving combustion efficiency.

Model predictive control–based compound control strategy for ducted rockets gas flow with anti-regulation suppression

Variable flow ducted rockets are excellent candidates for propulsion systems in hypersonic vehicles. However, the anti-regulation characteristics of the gas flow not only decrease the stability of ducted rockets but also create difficulties in the control system design. In this study, a model predictive control (MPC)–based compound control algorithm that suppresses the anti-regulation is proposed. First, a mathematical model involving a gas flow regulation system was developed to describe and analyse the anti-regulation. Next, a fused feedback variable that included both the gas flow and the gas generator pressure signal was constructed. An MPC control system with the fused feedback variable was also designed. Then, to accurately predict the gas flow and the anti-regulation, a nonlinear disturbance observer was constructed. To evaluate the performance of the controller, two scenarios that were regulated by a tracking differentiator proportional–integral–derivative controller and a model reference adaptive control controller were constructed and compared. The theoretical results demonstrate that the proposed method exhibited a superior performance compared to previous methods by overcoming the supposed trade-off between the response rate and the anti-regulation of gas flow systems.

Closed-Loop Fault Diagnosis of SDR Using Gap Metric, PCA, and Kalman’s Principle

The measurement of the pressure in the gas generator (GG) plays a decisive role in the closed-loop regulation of the gas flow in a solid ducted rocket ramjet (SDR). Therefore, the fault detection and isolation (FDI) of the pressure sensors is particularly significant. Especially in the GG with double pressure feedback, there are issues such as less available data and sensor fault transmission, which make the FDI more difficult. In this paper, for the GG with double pressure feedback, firstly, the “Consistency Index” was constructed based on the principal component analysis (PCA) algorithm, which made it easier to evaluate the consistency of the measured values. Secondly, based on Kalman’s principle, the redundancy information of the system was used to estimate the pressure in the GG. Finally, the gap metric between the measured pressure and the estimated pressure was employed to characterize the health of the sensors. Compared with the MWPCA algorithm, it was shown that our proposed algorithm was more accurate in fault diagnosis and could avoid the problem of missing alarm when two sensors had consistent faults, which would provide strong support for the safe operation of the SDR and could further promote its application in long-endurance aircrafts.

Hybrid Modeling Method of SDR Interstage Valve Based on Mechanism and Data-Driven

Variable-throat adjustment is the most practical flow regulation method of solid ducted rocket ramjet (SDR). The high-fidelity mathematical model of the interstage valve is the basis for realizing high-precision gas flow and thrust regulation. In this paper, the complex effect of gas was divided into load and throat deformation effect. The load was mainly determined by the clearance, friction torque, and pneumatic torque that the valve was subjected to during operation. And the throat deformation was determined primarily by the deposition and ablation of the valve faced in the gas. Therefore, we could divide the valve model into three parts: the servo motor model, the load characteristic model, and the deformation model of the actual acting throat (referred to as the throat). Given, we have designed a cold-air experiment program, using cold air to equalize the valve load. Furthermore, we analyzed its mechanism of action and established the load model using the experimental data and neural network. Finally, the deformation mechanism of the throat was investigated, and simultaneously, the deformation model was shown based on the flight test data. Compared with the traditional interstage valve model, the model established in this paper is closer to the actual working conditions, which is helpful to carry out the more comprehensive and practical ground simulation. It has essential reference value for further realizing the precise regulation of gas flow.

Adaptive Control System Design and Experiment Study of Gas Flow Regulation System for Variable Flow Ducted Rockets

Variable flow ducted rockets (VFDRs) are promising candidates for propulsion systems in hypersonic vehicles because of their inherent advantages, such as high specific impulse, low weight, and high speed. The control of gas flow is essential for optimal VFDRs performance. However, the characteristics of gas flow regulation systems, such as anti-regulation, non-linearity, and parameter variation, make it difficult to construct gas flow controllers. Aiming at the above problems, we propose a compound control strategy integrating a novel second-order fuzzy adaptive tracking differentiator (SOA-TD) and an intelligent proportional-integral controller based on adaptive neuro-fuzzy inference system (ANFIS). First, a mathematical model of a gas flow regulation system was developed to analyze the control characteristics of VFDRs. Next, an ANFIS-based proportional-integral controller to developed to respond to the system’s time-varying characteristics. In addition, a novel SOA-TD was constructed to optimize the “arrange transient process” of instructions, which effectively suppressed anti-regulation of the gas flow without increasing response time. Finally, a hardware in loop (HIL) simulation device for VFDRs was established, and serial HIL simulation tests were carried out to verify the validation of the controller. The HIL simulation results indicate that our strategy exhibited a superior performance compared to traditional controllers in terms of adaptability, ability to suppress anti-regulation, and robustness, which is hoped to fulfill VFDRs’ thrust control requirements for a wide range of altitudes and Mach numbers in future engineering applications.

Application of Gap Metric to LADRC Design in Multilinear Model of SDR

The solid ducted rocket ramjet (SDR) system faces many disturbances in the process of operation, and the linear active disturbance rejection controller (LADRC) has been widely used in engineering to solve such problems. However, the SDR also has strong nonlinearity, which brings a great challenge to the application of the LADRC in the gas flow regulation of the SDR. And the problem of fast adjustment of the “compensation factor” is one of the main difficulties in LADRC. In this paper, under the LADRC frame, the gas generator system’s closed-loop stability of the SDR was analyzed and the range of compensation factors had been calculated, and then, the “gap factor” was introduced and the “cross-iteration” method was used to quickly map out the “compensation factor” in the multilinear model controller based on the variation of the zero-point position of the system and the gap metric between adjacent set points. This greatly simplifies the parameter tuning process of the LADRC when it was applied to strongly nonlinear systems. Finally, through the comparison of simulation with adaptive PI controller and model-assisted LADRC (M-LADRC), the results have shown that the control method designed in this paper can obtain satisfactory performance and has a good engineering application prospect.

Virtual Free-Volume Revised Method and Adaptive Control for Solid Ducted Rockets

AbstractThe response difference between the classical mathematical model of a ducted rocket and a real experiment is compared in this study, and a new revised method is proposed based on the free v...

Adaptive Backlash Compensation Method Based on Touch State Observation for a Solid Ducted Rocket

A hysteresis model was built to describe the backlash of the flow regulator in a solid ducted rocket, and its influence on the engine control was also analyzed in this study. An adaptive backlash compensation method was proposed under two challenges: limited information and backlash state variation caused by the harsh environment in the gas generator. The touch state is designed and its observation is used to get the state of backlash, and a compensation control method using the existing information was carried out combined with the motion intention. This method greatly shortened the time during the transition and reduced the hysteresis effect on the control system. Furthermore, the compensation method is improved and acquires a self-learning ability, the compensation parameter changes adaptively during the process of flow regulation, and it is able to meet the challenge of an unknown and variable state of backlash. Finally, the validation of the compensation method was carried out with two simulations.

Performance Evaluation of Boron/Hydroxyl-Terminated Polybutadiene-Based Solid Fuels Containing Activated Charcoal

Numerous researches have been conducted on boron-based fuels and propellants due to the attractive theoretical energy potential of boron. However, sustained energy production during combustion of a boron-based fuel/propellant is still uncertain because efficient burning of boron is challenging. The present paper deals with various boron-loaded polymeric fuels, where activated charcoal (AC) is introduced as a burning enhancer for application in hybrid fuel ducted rocket. The solid fuel samples are tested with the help of a classical screening tool, which is known as opposed flow burner (OFB) or stagnation flow burner. Opposed flow burner is run with oxygen mass flux between 20 and . The fuel is evaluated on the basis of major performance parameters, such as regression rate, species formation during combustion using a spectrometer, residual active ingredients in postcombustion product using standard material characterization techniques, etc. The major outcome is evaluated from thermogravimetric analysis (TGA), where active residual boron is found to be around 6% only in postcombustion product for the fuel sample containing B and AC in equal proportion. Opposed flow burner helps in reducing the time and cost of research involving expensive fine fuel ingredients, and also facilitates to identify optimum fuel formulations for hybrid gas generator in ducted rocket applications.

An LES for the Effects of Particle Size and Velocity Ratio on the Flow Structure in the Secondary Combustor

The effects of flow structure and particle size in the secondary combustor of a ducted rocket on mixing are investigated using a two-phase large eddy simulation analysis. The corner recirculation zone and central recirculation area appear in the combustor. Vortexes are generated mainly by the axial inlet flow. The stretched eddies along the axial direction appear on the rear section of the combustor. The mainstream in the combustor shows a periodic oscillatory repetition with a frequency of about 1100 Hz. This motion is an essential factor in determining the distribution of droplets. When the size of the particles exceeds 10 micrometers ( $${\rm St}_{\rm k}$$ > 1), the particles do not penetrate the corner recirculation zone. The flow distribution according to the size of the particles does not differ significantly, but the distribution of particles can be very different, which is an important consideration for combustion.

Combustion of Nanoaluminum and Magnesium in Fuel‐Rich Propellants

AbstractFuel‐rich propellants are used in ducted rocket technology. The reduced content of oxidizer and the use of atmospheric air improve the gravimetric specific impulse far above the level of standard rocket motors, granting longer operating range. Peculiar requirements of ducted rockets favor either high or low‐pressure exponent propellants, depending on the pressurization strategy of the gas generator. The introduction of metal additives further improves the volumetric specific impulse and plays in favor of system compactness but combustion is difficult due to the intrinsic lack of oxidizer. Agglomeration of metal and low burning rates are observed. In this paper, the combustion of a set of fuel‐rich propellants with nanoaluminum and magnesium is investigated. The reported data set comprises ideal performance, burning rate, incipient agglomeration, and metal content of the agglomerates.

Load Characteristics and Modeling Methods for the Flow Regulator of a Solid Ducted Rocket

The load characteristics for the flow regulator of a solid ducted rocket are discussed in this paper. The mechanism and the influence factor of the load in the flow regulator were studied both theoretically and experimentally, and the system load and the working load were divided according to the mechanism. Additionally, the load influence on the working quality of the flow regulator and the ducted rocket were analyzed. System modeling of the flow regulator was carried out based on the working mechanism of the gas generator and the Stribeck friction model, and models of the actuator, gas generator, and load were built. Furthermore, considering the identification of the model parameters and the inapplicability of the conventional Stribeck identification method to the flow regulator, a new Stribeck model parameter identification method was proposed. A sine wave was used as the signal source, and specially designed devices were used to accomplish the switch of the working condition. Through the use of the genetic algorithm, the rotational inertia and parameters of the system load and working load were identified by stages. Finally, the validation of the identification method was carried out with a simulation.

Investigation of a Ducted Rocket with a Paraffin/Oxygen Hybrid Gas Generator

Analysis and testing of a ducted rocket employing a hybrid gas generator is presented. Over 150 successful static firing tests have been performed using a connected pipe facility supplying vitiated air to the ram combustor at stagnation temperature and pressure encountered in actual flights. The hybrid gas generator consisting of paraffin wax and gaseous oxygen operated unchoked, supplying hot, fuel-rich gases to the ram combustor for further burning with air. Empirical correlation of the fuel regression rate versus oxidizer mass flux in the gas generator has been displayed, and good overall combustion efficiency has been revealed. Thrust and fuel/air ratio control via the cold oxidizer flow rate into the gas generator is a major advantage of employing a hybrid gas generator. It was shown that such control could be done over broad ranges of flight altitude, flight Mach number, and thrust, while maintaining the ram pressure throughout the engine. This is in contrast to the thrust control of a ducted rocket employing a fuel-rich solid propellant. In that case, the process involves variation of the gas generator nozzle throat area by a mechanical device within the hot flow, implying a large increase of pressure in the gas generator combustor. Different scenarios of engine control and thrust modulation for variable flight conditions and requirements have been demonstrated.

Experimental Observation and Characterization of B−HTPB‐based Solid Fuel with Addition of Iron Particles for Hybrid Gas Generator in Ducted Rocket Applications

AbstractAn extensive diversity in the formulation of HTPB‐based solid fuels with metal additives has now been accepted in rocket and missile applications. Boron, which is a metalloid and often occurs in compound form has maintained its dominance mainly due to its significant theoretical energy potential. Therefore, it is preferred as the primary choice for rockets and solid fuel ramjet (SFRJ) or ducted rocket (SFDR). In this paper, performance enhancement of B−HTPB‐based solid fuel is assessed through a rapid screening tool which is called an opposed flow burner (OFB). Screened fuel combinations may be further evaluated in the gas generator of a ducted rocket (DR) to fulfill the intention of the present investigation. The baseline fuel (B−HTPB) is included with a metal fuel/catalyst (iron) with the expectation that it would enhance the combustion performance of boron. Boron−iron loading has been taken 10 % of the total sample weight (wt.) giving a wt. ratio of 90 : 10. Pure oxygen has been used as an oxidizer which impinges on the fuel surface where oxygen mass flux (Gox) varies between 20–57 kg/m2‐s. In this study, the research focuses mainly on regression rate and amount of residual active boron in the condensed combustion product (CCP) of the samples. Addition of iron to B−HTPB increases the regression rate by around 12 % and active boron content in CCP reduces significantly to 2.5 % compared to 16 % of the baseline fuel. In addition to that, the heating value of the fuel samples and identification of gas phase intermediate species (OH, CH, C2, BO, BO2) are the other vital investigations undertaken in the present study.

Effects of Ti and Mg particles on combustion characteristics of boron–HTPB-based solid fuels for hybrid gas generator in ducted rocket applications

In the field of propellants and explosives, boron has been considered as an attractive fuel due to its calibre to produce high energy density. Therefore, boron as a fuel has great prospect for application in solid fuel ducted rocket (SFDR) with hydroxyl-terminated polybutadiene (HTPB) which is widely used solid fuel. Experimentally, it has been observed that ignition and combustion of boron is problematic due to boron oxide (B2O3) coating around active boron which inhibits its ignition/combustion. In the present study, titanium and magnesium have been used with boron-HTPB based solid fuel as burning promoters. Various combustion parameters have been evaluated for nine different fuel compositions. An opposed flow burner (OFB) system has been used for fuel evaluation with the help of gaseous oxygen (GOX) impingement where oxygen mass flux (Gox) is maintained between 20 and 57 kg/m2-s. Favorable result has been observed for boron-magnesium combinations as compared to boron-titanium. Around 18% enhancement in regression rate has been achieved for boron-magnesium-HTPB sample (at max. Mg %) compared to that of boron-HTPB sample. In all the boron-based samples, greenish appearance has been observed during combustion which corresponds to BO2 emission. Spectroscopy analysis confirmed the same. Burning events have been analysed using color and high-speed camera. Condensed combustion products (ejected from the burning surface) have been analysed by various material characterization techniques such as: FE-SEM, EDX, XRD, and TGA. Through TGA, the active residual boron content in the condensed product of magnesium-based sample is found to be 33%. This value is found to be the lowest among the tested fuel compositions. The present investigation has been mainly focused on the formulation of new composition of boron-based solid fuel having high energy density.