Liquefying fuel regression rate modeling in hybrid propulsion

Lestrade, Jean-Yves; Anthoine, J.; Lavergne, G.

doi:10.1016/j.ast.2014.11.015

Cited by 29 publications

(18 citation statements)

References 6 publications

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“…The injection subsystem (3) deals with the axial and tangential oxidizer injectors. Both the total oxidizer mass flow rate and the effective swirl number have been defined previously by Equations (4) and (6) and these definitions are used for the simulations. The total oxidizer mass flow rate is the sum of the axial and tangential oxidizer mass flow rates, and the effective swirl number depends on the geometric swirl number.…”

Section: Simplified Modeling Of A-soft Engines For Simulationsmentioning

confidence: 99%

“…These fundamental issues have been investigated and solutions do exist depending on the mission and engine configuration. Although a low regression rate may be interesting for long duration and low thrust missions, a higher regression rate can be obtained by using liquefying fuels [3,4]. By increasing the residence time of propellants and improving their mixing, swirl oxidizer injection enhances the combustion efficiency making the hybrid engine competitive regarding that aspect [5,6].…”

Section: Introductionmentioning

confidence: 99%

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Theoretical Investigation on Feedback Control of Hybrid Rocket Engines

Messineo

Shimada

2019

Aerospace

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Despite the fact that hybrid propulsion offers significant benefits, it still suffers from some limitations such as the natural oxidizer to fuel ratio shift which induces variations of the engines’ performances while operating. To overcome that issue, Japan Aerospace Exploration Agency (JAXA) has been studying an innovative concept for several years based on the combination of controlled axial and radial oxidizer injections, called altering-intensity swirling-oxidizer-flow-type engine. This type of motor is theoretically capable of managing both the thrust and the oxidizer to fuel ratio independently and instantaneously by using a feedback control loop. To be effective, such engines would require in-flight instantaneous and precise thrust and an oxidizer to fuel ratio measurements as well as an adapted feedback control law. The purpose of this study is to investigate the effect of measurement errors on the engine control and to propose a regulation law suitable for these motors. Error propagation analysis and regulation law are developed from fundamental equations of hybrid motors and applied in a case where resistor-based sensors are used for fuel regression rate measurement. This study proves the theoretical feasibility of hybrid engines feedback control while providing some methods to design the engine and regression rate sensors depending on the mission requirements.

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Section: Simplified Modeling Of A-soft Engines For Simulationsmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Theoretical Investigation on Feedback Control of Hybrid Rocket Engines

Messineo

Shimada

2019

Aerospace

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“…It follows that _ m end ∕ _ m ini 0.998 ≈ 1. Supposing that the k Ross ratio is constant during the firing test finally leads to (13) where f end ∕f ini 0.93 and predicts a frequency decrease of 7% (30 Hz). This value is closer than the one predicted with Eq.…”

Section: Stability Analysismentioning

confidence: 99%

“…Liquefying fuels are investigated because they can provide higher regression rates and thrust than classical fuels [10][11][12]. Lestrade [13,14] has developed an experimental facility to validate a one-dimensional numerical code for the liquefying fuel regression rates evaluation in hybrid rocket motors. The motor is classically composed of an injector, a prechamber including an igniter, a combustion chamber, a postcombustion chamber, and a nozzle.…”

Section: Introductionmentioning

confidence: 99%

Vortex Shedding Influence on Hybrid Rocket Pressure Oscillations and Combustion Efficiency

Messineo

Lestrade

Hijlkema

et al. 2016

Journal of Propulsion and Power

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Numerical simulation is performed and compared with a firing test of an H 2 O 2 ∕polyethylene hybrid rocket engine equipped with a catalyzer. The simulation is performed assuming an axisymmetric flow, unsteady Reynolds-averaged Navier-Stokes equations are solved based on a single-phase flow. Stable pressure oscillations are observed during the firing test and could be explained by the simulation with the formation of large-scale vortices at the end of the fuel grain and in the postcombustion chamber. These vortices transport unburnt fuel to the nozzle and play a major role in the mixing process in the postchamber. Two semi-empirical approaches are applied and provide a good order of magnitude for the frequency and frequency shift during the firing test. Another simulation has been performed with an added oxidizer postchamber injection. It revealed that small-scale vortices along the fuel grain in the baseline case are probably generated by the oscillatory motion of the flow due to pressure fluctuations in the engine. Nomenclature C μ = turbulence model constant c = speed of sound, m · s −1 c th = theoretical characteristic velocity, m · s −1 D = fuel port diameter, m D t = nozzle throat diameter, m f = frequency, Hz k = turbulence kinetic energy, m 2 · s −2 k Ross = ratio of vortex transport velocity to freestream velocity L = total combustion chamber length, m L f = fuel grain length, m L pre ch = length of the prechamber, m L Ross = distance between shear layer initiation and impingement point, m l = exhaust nozzle length, m M = Mach number m = stage number (number of coherent structures) _ m = mass flow rate, kg · s −1 O∕F= oxidizer-to-fuel mixture ratio p = pressure, Pa Q = Q criterion, rad 2 · s −2 R = specific gas constant, J · kg −1 · K −1 R f = fuel port radius, m r = distance from wall, m S = symmetric part of velocity gradient tensor S D = Strouhal number based on end-port diameter T = temperature, K T Ross = period, s U = freestream velocity, m · s −1 U vortex = vortex transport velocity, m · s −1 V = combustion port total volume, m 3 V x = axial velocity, m · s −1 x = axial position, m Y = mass fraction y = radial position, where y equal to zero is the axis, m α = empirical constant Δt = correction factor, s η c = combustion efficiency κ = von Kármán constant ρ = density, kg · m −3 τ bl = boundary-layer delay time, s ψ = function of isentropic coefficient Ω = skew-symmetric part of velocity gradient tensor ω = turbulence specific dissipation rate, s −1 Subscripts 1L = first longitudinal acoustic mode end = end of firing test H = Helmholtz mode hy = hybrid intrinsic instability ini = beginning of firing test ox = oxidizer

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“…[6][7][8] Finally, appropriate models are available for predicting the regression rate of classical or liquefying fuels. 9 Nammo Raufoss has also developed a know-how of more than 10 years on hybrid propulsion. Nammo initiated in 2003 an in-house technology program to study hybrid rocket technology and made its first static 10,11 More recently, Nammo has worked on the development of a hybrid rocket engine for a lander application 12 and on the large thrust hybrid propulsion demonstrator within the ESA FLPP program.…”

Section: Introductionmentioning

confidence: 98%

Experimental Demonstration of the Vacuum Specific Impulse of a Hybrid Rocket Engine

Anthoine¹,

Lestrade²,

Verberne³

et al. 2014

50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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As part of the cooperation between Onera and Nammo Raufoss on hybrid propulsion, a small hybrid rocket engine was tested in the Onera vacuum test facility to demonstrate the experimentally achievable vacuum specific impulse. The Onera hybrid rocket engine, fitted with regression rate sensors, was modified to accommodate the catalyst bed and a vortex injector, both provided by Nammo Raufoss. The demonstrated combustion efficiency was above 98% and the specific impulse was experimentally proven to be above 270 s for a conical nozzle with an expansion ratio of 50.

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Liquefying fuel regression rate modeling in hybrid propulsion

Cited by 29 publications

References 6 publications

Theoretical Investigation on Feedback Control of Hybrid Rocket Engines

Theoretical Investigation on Feedback Control of Hybrid Rocket Engines

Vortex Shedding Influence on Hybrid Rocket Pressure Oscillations and Combustion Efficiency

Experimental Demonstration of the Vacuum Specific Impulse of a Hybrid Rocket Engine

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