Performance study of a water ramjet engine

Huang, Liya; Xia, Zhixun; Hu, Jie; Zhu, QianWen

doi:10.1007/s11431-010-4242-7

Cited by 37 publications

(22 citation statements)

References 14 publications

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“…Thrustmeasuring components include leaf springs, a flexible universal joint and a thrust meter. 15) Through a flexible universal joint, the movable frame is connected to a thrust meter (model BLR-12 made by Hefei Zhongya Sensor Limited Company) with a maximum range of 1000 N.…”

Section: Experiments Apparatusmentioning

confidence: 99%

Experimental Investigation and Numerical Simulation of Secondary Combustor Flow in Boron-Based Propellant Ducted Rocket

Xia

Wang

et al. 2013

Trans. Japan Soc. Aero. S Sci.

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This paper establishes a two-phase reacting flow model in the secondary combustor of a ducted rocket. The threedimensional Favre-averaged compressible turbulent N-S equations are used as the governing equations of the reacting flow, the improved k $ " two-equation turbulence model is used to simulate the turbulent flow, and the eddy break up model is used to simulate the gas combustion. The particle-phase solution is obtained using a well-established boron particle ignition and combustion model. Boron particles are ejected from the exit of the gas generator into a secondary combustor and their trajectories are traced through the reacting flowfield using discrete phase models. The secondary combustion in a cylindrical combustor for the ducted rocket is investigated preliminarily with tests conducted on a connected-pipe ramjet test bed. During the test, boron-based fuel-rich HTPB propellant is used. The effect of factors, such as air/fuel ratio, velocity of air injection and dome height on the performance of the combustor and the engine is examined. The work described in this paper represents an attempt to direct the design of the secondary chamber in order to increase combustion efficiency. The experimental results show that the reacting flow model established in this paper is correct.

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Section: Experiments Apparatusmentioning

confidence: 99%

Experimental Investigation and Numerical Simulation of Secondary Combustor Flow in Boron-Based Propellant Ducted Rocket

Xia

Wang

et al. 2013

Trans. Japan Soc. Aero. S Sci.

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“…In this study, the characteristic velocity of the hybrid rocket motor was adopted to calculate its combustion efficiency. An experimental characteristic velocity and combustion efficiency were given by 18) where p c was the aft mixing chamber pressure, A t was the throat area of the hybrid rocket motor nozzle, q ox and q f were the mass flow rate of the gaseous oxygen and fuel, c Ã and c Ã th indicated the experimental and theoretical characteristic velocities of the combustion products of hybrid combustion chamber, and c Ã was the combustion efficiency of the hybrid combustion chamber.…”

Section: Combustion Efficiencymentioning

confidence: 99%

Experimental Investigation of Fuel Regression Rate in Hydroxyl-terminated-polybutadiene Gaseous-oxygen Hybrid Rocket Motors

Zhang

Zhi-bin

et al. 2012

Trans. Japan Soc. Aero. S Sci.

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The regression rate enhancement of hydroxyl-terminated-polybutadiene (HTPB) gaseous oxygen (GOX) hybrid rocket motors was investigated experimentally in a ground testing system. The effects of the addition of nano-sized aluminum (Al) or alloyed aluminum-magnesium (Al-Mg) powder to the fuel composition, mass flux of oxygen, and the oxidizer-fuel ratio on the fuel regression rate are presented. The experiments show that multiple-start and shutdown can be realized using gaseous oxygen/RP-1 pre-mixed torch designed in the present study. The fuel surface remained hot enough during the shutdown interval to reestablish combustion if the time of motor shutdown is less than 2 s. The deduced spatially averaged instantaneous regression rate and spatially and time-averaged regression rate fit well in the trend with time. The former way to calculate the regression rate can more precisely describe the fuel combustion process. As the oxidizer-fuel ratio increased during the test, the hybrid rocket motors didn't work at the optimal oxidizer-fuel ratio, which can induce a loss of specific impulse. The addition of alloyed Al-Mg or nano-sized Al to the HTPB fuel can enhance the regression rate. The effect due to the latter is the most significant.

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“…As for the hydroreactive fuel of the engine, the high heat release of Al∕H 2 O and Mg∕H 2 O reactions have made aluminum and magnesium powders the most promising high-energy additives [4,5].…”

Section: Introductionmentioning

confidence: 99%

Experimental Investigation on Combustion of High-Metal Magnesium-Based Hydroreactive Fuels

Chao

Xia

et al. 2013

Journal of Propulsion and Power

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The experimental study in this paper focused on the combustion wave temperature measurement and the qualitative observation of phenomena involved in the combustion of a high-metal magnesium-based hydroreactive fuel strand, using a pressure-regulated test combustor. An experimental system was designed and experiments were carried out in both argon and water vapor atmosphere. Two types of hydroreactive fuel were investigated: one containing 60% magnesium particles and the other one containing 73% magnesium particles. Thermal-wave structures of the combustion wave of the hydroreactive fuel were measured by an imbedded type-K fine-wire thermocouple. The surfaces of fuel samples quenched by rapid depressurization were examined in a scanning electron microscope. The chemical composition on the burned surface of the hydroreactive fuel strand was determined by Xray diffraction analysis. The experimental phenomena implies that magnesium particles in the 60% magnesium fuel are ejected into the gas phase and combust, whereas magnesium particles in the 73% magnesium fuel stay on the burning surface and combust. Using these experimental results, the effect of the amounts of metal additives on the combustion of the high-metal magnesium-based hydroreactive fuel was postulated and the physical combustion model was proposed. Nomenclature c p = specific heat, J∕kg · K −1 r = burning rate, mm∕s T = temperature, K t ig = igniting time, s Q = rate of heat generation, J∕kg λ = thermal conductivity, W∕m · K −1 ρ = density, kg∕m 3Subscripts c = solid phase f = combustion flame g = gas phase s = burning surface of strand 0 = initial

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Performance study of a water ramjet engine

Cited by 37 publications

References 14 publications

Experimental Investigation and Numerical Simulation of Secondary Combustor Flow in Boron-Based Propellant Ducted Rocket

Experimental Investigation and Numerical Simulation of Secondary Combustor Flow in Boron-Based Propellant Ducted Rocket

Experimental Investigation of Fuel Regression Rate in Hydroxyl-terminated-polybutadiene Gaseous-oxygen Hybrid Rocket Motors

Experimental Investigation on Combustion of High-Metal Magnesium-Based Hydroreactive Fuels

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