Fuel-rich Propellant Research Articles

Combustion Characteristics Experimental Study of Solid Hydrocarbon Propellant for Air-Turbo Rocket

The combustion characteristics of propellants heavily affect the performance of solid-propellant air-turbo rockets (SP-ATRs). In this study, an SP-ATR experimental combustion system was developed, and experimental studies were conducted on the primary and secondary combustion characteristics of two solid hydrocarbon fuel propellants. The results show that the solid hydrocarbon fuel-rich propellant, which is an ideal propellant for SP-ATRs, can play the dual role of driving the turbines and executing secondary combustion. The studies on the two candidate propellants show that CH01 has higher energy characteristics, and CH02 exhibits better performance in terms of the interior ballistic stability and ejection efficiency. For cold inflow conditions, when the fuel gas flow rate is at the optimum equivalence ratio, there are difficulties in secondary combustion. A combustion experiment with a variable gas flow rate was designed, and the upper limit of the excess air coefficient was obtained under the condition of cold inflow.

Combustion Characteristics of Ramjet Fuel Grains with Boron and Aluminum Additives

An experimental investigation of the effect of boron (B) and aluminum (Al) additives to fuel-rich propellant for a solid fuel ramjet application was performed. In this research, depending on the composition of metal particles, three types of fuel grains with 15 wt % ammonium perchlorate particles were prepared and tested. To provide a short ignition delay, an ignition support material consisting of and a composite propellant were coated on the fuel grain. For the test, an oxidant gas having a controlled temperature and pressure and an oxygen composition close to that of the air in a ramjet combustor was supplied using an ethanol-blended hydrogen peroxide () gas generator. Air simulant gas was supplied with an average mass flow rate near and a Mach number of 0.21 in the fuel grain port. Through the test, the noticeable effect of B particles on combustion was confirmed by a high regression rate over and combustion efficiency over 80%. Al particles were not effective additives individually; however, Al particles contributed to the ignition of B. In addition, metal oxide particle deposits on the nozzle confirmed the existence of a two-phase flow.

Concept and performance study of turbocharged solid propellant ramjet

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

The effect of pressure and oxygen concentration on the ignition and combustion of aluminum–magnesium fuel-rich propellant

In order to better understand how future candidate solid propellants may affect combustion characteristics in ramjet motor, the aluminum–magnesium fuel-rich propellant was tested on a sealed high-pressure laser ignition platform. The laser ignition experimental device was designed and built, then the ignition and combustion characteristics of aluminum–magnesium fuel-rich propellant under different pressures and oxygen concentrations in environment gas were studied. The comprehensive impacts of pressure and oxygen concentration on ignition delay time and burning rate were appraised. The results displayed that the ignition delay time was decreased with the increase of pressure and oxygen concentration. The effect of pressure on ignition delay time was weaker than oxygen concentration. The impact of oxygen concentration on ignition was very complex. When the oxygen concentration of environment gas was less than that of the oxidizing gas from the propellant pyrolysis products, the ignition gas phase reaction occurred in the dispersion region of the solid propellant pyrolysis product, the initial flame was close to the propellant surface. In contrast, when the oxygen concentration was larger than the pyrolysis oxidizing gas, the ignition gas phase reaction occurred in the diffusion zone of the propellant pyrolysis products and the environment gas, the initial flame appeared far from the propellant surface. The burning rate of aluminum–magnesium fuel-rich propellant increased with the increase of pressure and oxygen concentration, where their effect on burning rate is accord with B-number theory. Comparing the experimental measurements with an analytical model, the results showed good consistency. The function of oxygen concentration and pressure could explain the effect on burning rate and ignition delay time, because pressure and oxygen concentration were favorable for the heat transfer, radiation and convection from the flame. The function of the combined influences of pressure and oxygen concentration on burning rate was established by a power law fitting. The good fitting is consistent with the experimental measurements, and shows that the scaling is valid for a full range of pressure and oxygen concentration tests to predict burning rate.

Thermal behavior of Al/Zr/KClO4 pyrotechnic compositions at high temperature

A pyrotechnic composition of aluminum, zirconium, and potassium perchlorate was prepared to obtain suitable safety and sensitivity. The thermal behavior, microstructure, and composition of the pyrotechnic compositions heated from 300°C to 600°C were investigated using differential scanning calorimetry (DSC), thermogravimetry (TG), scanning electron microscopy (SEM), and energy-dispersive X-ray spectroscopy (EDX) to evaluate the properties of the pyrotechnics after the high-temperature process, such as the 1A/1W no-fire test. When the heating temperature increased, the sharp exothermic peak near 530°C decreased, corresponding to the reaction of metal powders with potassium perchlorate. Gradual oxidizing reaction between Zr and KClO4 can occur in solid stage when the temperature exceeds 400°C. With the increase of the heating temperature from 400°C to 600°C, the mass losses in TG curves decreased, and melted particles of KClO4 appeared gradually. The Al/Zr/KClO4 pyrotechnic composition was safe in the of 1A/1W no-fire test. Hence, Al/Zr/KClO4 can be used in flash pyrotechnics, fireworks, and fuel-rich propellants.

램젯 연소실 조건에서 점화보조제가 도포된 Fuel-rich 추진제의 점화

램젯 연소실 조건에서 점화보조제가 도포된 Fuel-rich 추진제의 점화 실험을 수행하여 점화 지연과 연소 유지를 확인하였다. 연료 그레인은 HTPB에 AP 파우더 15 wt.%, Al 파우더 5 wt.%가 혼합된 형태로 구성되어 있다. 한편 연료 그레인에 <TEX>$NC/BKNO_3$</TEX>와 Composite 추진제로 이루어진 점화보조제를 도포하여 빠른 점화지연이 나타나도록 하였다. 에탄올 블렌딩 과산화수소 가스발생기를 통해 램젯 연소실의 공기와 가깝도록 온도, 압력, 산소 조성을 조절한 산화제 가스를 유속 <TEX>$200kg/m^2s$</TEX>으로 공급하였다. 실험 결과, 점화보조제가 잘 작동하여 연료그레인에서 0.6초의 점화 지연시간을 파악하였고 연소 중에는 화염이 유지되는 것을 확인하였다. Ignition test of the fuel-rich propellant coated with ignition support material in the ramjet combustor condition was conducted. Ignition delay and flame holding was measured. Fuel grain consist of HTPB mixed with AP particle 15 wt.%, Al particle 5 wt.%. To cause the short ignition delay, ignition support consist of <TEX>$NC/BKNO_3$</TEX> and composite propellant was coated to the fuel grain. Ethanol blended <TEX>$H_2O_2$</TEX> gas generator control the temperature, pressure, <TEX>$O_2$</TEX> concentration in the oxidizer gas in the air. Gas is supplied with mass flux of <TEX>$200kg/m^2s$</TEX>. Through the test ignition support operated well and ignition delay of 0.6 second and the Flame was sustained.

Ignition and Combustion Performance of the Primary Condensed-phase Combustion Products from Boron-based Fuel-rich Propellants

Ignition and Combustion Performance of the Primary Condensed-phase Combustion Products from Boron-based Fuel-rich Propellants

Rate constants determination for reactions of B2O3 with H2O and HCl at high temperature using ab initio calculations

Primary combustion of boron-based fuel-rich propellants produces large amounts of HCl(g) and H2O(g) which may react with B2O3(g) during the secondary combustion at high temperature. The reaction paths of the two reactions were expedited with Gaussian-09 program at CCSD(T)/aug-cc-pvtz//M062X/aug-cc-pvtz level, and the rate constants and Arrhenius parameters were estimated by using conventional transition state theory, variational transition state theory and variational transition state theory with interpolated single-point energies. The calculation results show that rate constant of the two reactions is in the range of 6.17×10−16 to 3.57×10−15cm3molec−1s−1 and 1.58×10−16 to 1.47×10−15cm3molec−1s−1 at temperature 2500–3500K with activation energy of 118.01–136.29kJ/mol and 155.12–168.28kJ/mol, respectively.

Attaining Hypersonic Flight with Aluminum-Based Fuel-Rich Propellant

In this paper, experimental results of an aluminized fuel-rich propellant having energetics similar to boron-based fuel-rich propellant and having no residue after combustion are discussed. Using this aluminum-based fuel-rich propellant, which has higher density than the current fuels used for hypersonic flight, this paper looks at enhancing the Mach number envelope of a ramjet. Nondimensional thrust and drag calculations were carried out to identify the air–fuel ratio required for a Mach 6 flight, with different intakes accounting for different pressure recovery. Calculations were carried out with three different intakes, and it was observed that, even with the lowest pressure recovery of 9% at Mach 6, a hypersonic mission was possible. Values for various efficiencies were chosen with great care, and a perturbation analysis was carried out to make the calculations more robust. The specific impulse obtained with the aluminum-based fuel-rich propellant in ramjet mode was 515 s at a 25 km altitude. Comparison of an aluminum-based fuel-rich propellant in ramjet mode was made with a kerosene and hydrogen fueled scramjet from the specific impulse data available in the literature. The specific impulse of the aluminum-based fuel-rich propellant was nearly an order of magnitude higher as compared to a hydrogen fueled scramjet when adjusted to the density of hydrogen. Upon considering a system size, it was also shown that the aluminum-based fuel-rich propellant could meet the burn rate requirement for a Mach 6 flight.

Effect of Different Coating Materials on Boron in Fuel Rich Propellant Composition Using Hydroxyl Terminated Polybutadiene as a Binder

Effect of Different Coating Materials on Boron in Fuel Rich Propellant Composition Using Hydroxyl Terminated Polybutadiene as a Binder

Evaluation of Epoxy Terminated Polybutadiene in Ammonium Perchlorate Coated Boron Based Fuel Rich Propellant Formulations

Evaluation of Epoxy Terminated Polybutadiene in Ammonium Perchlorate Coated Boron Based Fuel Rich Propellant Formulations

3,6-Dinitropyrazolo[4,3-c]pyrazole-Based Multipurpose Energetic Materials through Versatile N-Functionalization Strategies.

A family of 3,6-dinitropyrazolo[4,3-c]pyrazole-based energetic compounds was synthesized by using versatile N-functionalization strategies. Subsequently, nine ionic derivatives of the N,N'-(3,6-dinitropyrazolo[4,3-c]pyrazole-1,4-diyl)dinitramidate anion were prepared by acid-base reactions and fully characterized by infrared, multinuclear NMR spectra, and elemental analysis. The structures of four of these compounds were further confirmed by single-crystal X-ray diffraction. Based on their different physical and detonation properties, these compounds exhibit promising potential as modern energetic materials and can be variously classified as green primary explosives, high-performance secondary explosives, fuel-rich propellants, and propellant oxidizers.

Factors Affecting the Primary Combustion Products of Boron-Based Fuel-Rich Propellants

Boron-based fuel-rich propellants with different formulation were prepared, and the condensed-phase products under different chamber pressures were collected in this study. The composition of boride (elemental boron, boron oxide, boron carbide, and boron nitride) in the condensed-phase products was measured by chemical analysis and the particle size distribution was determined using a laser particle size analyzer. Moreover, the burning rate and combustion temperature of the propellants were measured using a strand burner. The experimental results show that both the propellant burning rate and combustion temperature played an important role in the production of boride. There might be large amounts of boron carbide in the condensed-phase products which appear to have a negative effect on the combustion efficiency of the propellants studied in this work. The production of boron carbide could be reduced and combustion efficiency was improved by lowering the combustion temperature while increasing the burning rate. This study also shows that reducing the chamber pressure also causes a decrease in production of boron carbide.

Effect of solid oxidizers on the thermal oxidation and combustion performance of amorphous boron

Amorphous boron is usually employed as the most important fuel of boron-based fuel-rich propellants, and NH4ClO4 (AP), cyclotetramethylenetetranitramine (HMX), KClO4 and KNO3 are the solid oxidizers of most used in solid propellants. The mixtures of boron and different solid oxidizers with mass ratio 1:1 were prepared in this paper, and the effect of these oxidizers on the thermal oxidation and combustion performance of amorphous boron was studied by simultaneous thermogravimetry–differential scanning calorimeter–Fourier transform infrared spectroscopy and CO2 laser ignition experiments. The experimental results show that the main reactions during the heating process of B/AP and B/HMX samples are the decomposition of oxidizers, and the decomposition process of oxidizers rather than the decomposition temperature is affected by amorphous boron; boron could react with KClO4 and KNO3 violently with the release of large amounts of heat, and then both of the oxidizers, especially KClO4, have positive effect on the oxidation and combustion performance of amorphous boron.

Experimental research on aluminium–magnesium fuel-rich propellant under high strain rates

The aluminium–magnesium fuel-rich propellant is one of the classified propellant used in the solid rocket ramjet. The weapons with the rocket ramjet are under high overload by launching, the mechanical properties of the propellant should be studied under high strain rate. The uniaxial compressive mechanical curves of the aluminium–magnesium fuel-rich propellant under high strain rates have been obtained with the modified split Hopkinson pressure bar in this paper. Four groups of stress–strain curves were obtained at different strain rates of 2600, 3300, 4500 and 5100 s−1. The experimental results show that the mechanical properties of the aluminium–magnesium fuel-rich propellant are rate-dependent, and the Mg–Al fuel-rich propellant has the mechanical properties of viscoelastic material. With the increase of strain rate, the modulus of the aluminium–magnesium fuel-rich propellant obviously increases. The underlying mechanism of the significant rate sensitivity was investigated by adopting the scanning electron microscope from a micro-structure point of view.

Ignition and combustion characteristics of compound of magnesium and boron

Compound of magnesium and boron (MB) is promising to be the ideal substitute of amorphous boron which is usually used as the raw material of boron-based fuel-rich propellants. In this study, the physical characteristics of MB and amorphous boron were studied by the scanning electron microscope, X-ray diffraction and X-ray photoelectron spectroscopy. The thermal reaction characteristics and the ignition and combustion characteristics were investigated through TG/DSC experiments and laser ignition experiments. The experimental results show that the MB particle is much more regular than amorphous boron, which favors for the preparation of boron-based fuel-rich propellants. Magnesium exists in the form of elementary substance, and boron oxide is produced during the preparation of MB which results in the longer ignition delay time of MB. The content of magnesium and the pressure have effects on the MB combustion performance. Although the combustion of magnesium can provide much heat for the combustion of boron, MB with moderate content of magnesium shows the best combustion performance. On the contrast, the ignition delay time of MB is independent on the content of magnesium and the pressure.

Energy release properties of amorphous boron and boron-based propellant primary combustion products

The microstructure of amorphous boron and the primary combustion products of boron-based fuel-rich propellant (hereafter referred to as primary combustion products) was analyzed by scanning electron microscope. Composition analysis of the primary combustion products was carried out by X-ray diffraction and X-ray photoelectron spectroscopy. The energy release properties of amorphous boron and the primary combustion products were comparatively studied by laser ignition experimental system and thermogravimetry–differential scanning calorimetry. The primary combustion products contain B, C, Mg, Al, B4C, B13C2, BN, B2O3, NH4Cl, H2O, and so on. The energy release properties of primary combustion products are different from amorphous boron, significantly. The full-time spectral intensity of primary combustion products at a wavelength of 580nm is ~2% lower than that of amorphous boron. The maximum spectral intensity of the former at full wave is ~5% higher than that of the latter. The ignition delay time of primary combustion products is ~150ms shorter than that of amorphous boron, and the self-sustaining combustion time of the former is ~200ms longer than that of the latter. The thermal oxidation process of amorphous boron involves water evaporation (weight loss) and boron oxidation (weight gain). The thermal oxidation process of primary combustion products involves two additional steps: NH4Cl decomposition (weight loss) and carbon oxidation (weight loss). CL-20 shows better combustion-supporting effect than KClO4 in both the laser ignition experiments and the thermal oxidation experiments.

Chemical analysis of primary combustion products of boron-based fuel-rich propellants

The content of solid primary products of boron-based fuel-rich propellants could be determined quantitatively through the analysis procedure.

Thermal decomposition of HTPB/AP and HTPB/HMX mixtures with low content of oxidizer

In this paper, hydroxyl-terminated polybutadiene (HTPB)/ammonium perchlorate (AP) and HTPB/cyclotetramethylene tetranitramine (HMX) mixtures with low content of oxidizer were prepared, and the thermal decomposition process of these mixtures was investigated with TG-FTIR and Raman spectrum. The experimental results indicated that during the thermal decomposition process under an inert atmosphere, HTPB and the two oxidizers could both decompose into gaseous products completely, while the mixture of HTPB and the two oxidizers produced some solid residue with carbon as the main ingredient; unlike the thermal decomposition of HTPB under argon atmosphere, HTPB cannot decompose completely under air or oxygen atmosphere, and the more the concentration of oxygen, the more the solid residue produced. Decreasing the content of oxidizer reacting with HTPB may improve the combustion performance of fuel-rich propellants by increasing the combustion temperature and the percentage of gaseous products. The results of this paper can provide the useful information for improving the combustion performance of fuel-rich propellants.

금속연료인 과립화붕소의 입도에 따른 연료과농 고체 추진제 특성 연구

공기흡입형 추진기관의 가스발생기에 사용하기 위한 연료과농(Fuel-rich) 고체 추진제의 특성에 대한 연구를 수행하였다. 일반적인 고체 추진제는 평균 60%이상의 산화제를 포함하는 데, 연료과농 고체 추진제를 개발히기 위해 산화제 함량을 30%내외로 낮추고 매우 높은 부피당 열량을 가지는 비정질 붕소를 입자화(Bead)하여 금속연료 함량을 증가시켜 고체 추진제의 제조 공정성 및 연소속도 특성을 확인하였다. 과립화붕소의 입도가 작을수록 추진제 제조 공정에서 초기 점도가 높아지고, 입도가 클수록 연소속도 및 압력지수가 증가하는 것을 확인할 수 있었다. A study of gas generator Fuel-Rich propellant for air-breathing propulsion system was performed in this paper. General solid propellant comprises a mean of 60% or more oxidizing agents. but, to develop the fuel-rich solid propellant increased the content of the metal fuel and reduced the content of the oxidizing agents by approximately 30%. Very high amount of heat per volume of fuel into the metal having the Boron was used. Amorphous Boron Powder was applied to propellant as beads type and it allowed to design more amount of metal fuel in the fuel-rich propellant. And the Combustion characteristics and properties of fuel-rich solid propellant according to the Boron-bead sizes were confirmed.