Multiphysics simulations of rocket engine combustion

Chen, Yen Sen; Chou, T. H.; Gu, B. R.; Wu, Jong-Shinn; Wu, B. H.; Lian, Y. Y.; Yang, Lijun

doi:10.1016/j.compfluid.2010.09.010

Cited by 44 publications

(44 citation statements)

References 9 publications

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“…For the closure of the above governing equations, an efficient method for treating the real-fluid equations of state and fluid properties is employed herein for liquid propellant combustion flows [10]. The convection terms of the governing equations are discretized with a second-order upwind scheme.…”

Section: Numerical Approachmentioning

confidence: 99%

“…Very few have attempted to model complicated reactive flow phenomena of a realistic hybrid propulsion systems 8-10 , they employ an energy-balanced surface decomposition model, assuming that HTPB decomposes primarily into C 4 H 6 and C 2 H 4 , starting around 800K, with some tuned rate constants for a set of reduced reaction mechanisms. Chen et al [10] proposes a mixing enhancement design in the combustion chamber with a comprehensive real-fluid model for the N 2 O injection. The overall efficiency and performance enhancement of such system has been validated with experimental data.…”

Section: Introductionmentioning

confidence: 99%

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N2O-HTPB Hybrid Rocket Combustion Modeling with Mixing Enhancement Designs

Chen

Lai

Chou

et al. 2013

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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One main physical feature of hybrid rocket combustion is its diffusion flame structure that requires excessively long solid grain port which often leads to undesirable large slenderness of a rocket configuration. The diffusion flame also results in generally low combustion efficiency of hybrid rockets. Some remedial designs have used liquefying solid grain, such as paraffin, or mixing enhancement mechanisms to boost the overall combustion efficiency. Thus, shortened combustion chamber can be used to deliver reasonable thrust performance of hybrid rockets. In addition to the study of multi-stage mixing enhancer effects, a compact hybrid rocket motor design concept is also proposed in the present study to provide better form factors for hybrid rocket engine designs. This design concept features in vertical-flow structures such that greatly improved combustion efficiency is obtained. A 3D computational model with finite-rate chemistry and radiative heat transfer capabilities is employed to assess the mixing effectiveness and combustion efficiency of the new design concept. The present computational model is validated for a wide range of rocket propulsion design problems, including a single-port hybrid rocket motor with and without using a mixing enhancement mechanism. The internal ballistics and flame structures in the hybrid rocket engine with mixing enhancement designs are analyzed. Nomenclature C 1 ,C 2 ,C 3 ,C  = turbulence modeling constants, 1.15, 1.9, 0.25, and 0.09. C p = heat capacity D = diffusivity F yz, F y, F z = integrated force, and component forces in the lateral direction Joint Propulsion Conferences 2 H = total enthalpy K = thermal conductivity or stiffness k = turbulent kinetic energy Q = heat flux T = temperature t = time, sec u = mean velocities V 2 =  u 2 x = Cartesian coordinates or nondimensional distance  = species mass fraction  = turbulent kinetic energy dissipation rate μ = viscosity μ t = turbulent eddy viscosity (=C  k 2 /) Π = turbulent kinetic energy production ρ = density  = turbulence modeling constants, 0.9, 0.9, 0.89, and 1.15 for Eqs. (2), (4-6). τ = shear stress ω = chemical species production rate or natural frequency Subscripts r = radiation t = turbulent flow

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Section: Numerical Approachmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

N2O-HTPB Hybrid Rocket Combustion Modeling with Mixing Enhancement Designs

Chen

Lai

Chou

et al. 2013

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference

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“…considered 13 species and 17 reactions for both combustion and pyrolysis [12], which seems to be the maximum mechanism size used in literature to the authors'…”

Section: Introductionmentioning

confidence: 99%

Detailed kinetic computations and experiments for the choice of a fuel–oxidiser couple for hybrid propulsion

Gascoin

Gillard

Mangeot

et al. 2012

Journal of Analytical and Applied Pyrolysis

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The choice of a solid reducer for hybrid propulsion is generally based on the quantity of gaseous combustible it can produce (expressed indirectly by the regression rate). For this reason, the studies focus on the use of additives or on the design of grain while the kinetic aspect is rarely of interest despite the chemistry drives the phenomena (chemical induction delay, heat absorption, chemical composition). One-step mechanisms are first considered in this paper to quantify the effect of operating conditions on high density polyethylene -HDPE-, polymethylmethacrylate -PMMA-and hydroxyl termination polybutadiene -HTPB-. Then the chemical composition of pyrolysis products is determined for a large range of operating conditions with highly detailed mechanism for HDPE (1014 species, 7541 reactions). The heating rate applied to the reducer is investigated (from 1 K s -1 to 10 7 K s -1 ). Ethylene is found to be the major pyrolysis product. The timescale found over 1250 K and 11.11 bar is in agreement with the requirements of hybrid propulsion.

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“…Unfortunately, even until now, research in developing the hybrid N 2 O-HTPB propulsion system still strongly depends on experimental trials-and-errors, which are time-consuming and expensive. Very few 10,11,12 have attempted to model complicated reactive flow phenomena of a realistic hybrid propulsion system, in which they employ an energy-balanced surface decomposition model with the assumption that decomposition of HTPB into C 4 H 6 (Iso-butadiene) starts at 820 K, and modeled the combustion process using some reduced reaction mechanisms with tuned rate constants.…”

Section: Introductionmentioning

confidence: 99%

“…As demonstrated in Ref. 12, the real-fluid effects with finite-rate chemistry can affect the overall flow structure in the combustion chamber, especially near the injectors, and affect the combustion processes and heat transfer characteristics, which is the key to properly predict the regression rates of the solid grain. Thus, the development of this comprehensive numerical model is continued in this research to study the flow physics of an N 2 O-HTPB hybrid rocket system.…”

Section: Introductionmentioning

confidence: 99%

Multiphysics Computational Modeling of Hybrid Rocket Combustion

Chen

Lian

et al. 2011

47th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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Recently, the hybrid rocket propulsion has become attractive to the research community and has developed the trend to become an alternative to the conventional liquid and solid rockets. Among available hybrid systems, the N 2 O (Nitrous Oxide)-HTPB (Hydroxyl-Terminated PolyButadiene) hybrid propulsion represents the simplest but sufficiently efficient design. To date, research in developing hybrid N 2 O-HTPB propulsion system, despite some available fuel regression rate correlations, still strongly depends on experimental trials-and-errors, which are time-consuming and expensive. Thus, detailed understanding of the fundamental combustion processes that are involved in the N 2 O-HTPB propulsion system can greatly impact the research community in this field. This may further facilitate the successful modeling of the combustion processes and help improving the design of N 2 O-HTPB propulsion system in the future. A comprehensive numerical model with real-fluid properties and finite-rate chemistry is developed in this research to predict the combustion flowfield inside a N 2 O-HTPB hybrid rocket system. The effects of a mixing enhancer design are demonstrated experimentally and numerically in boosting the overall thrust performance of a hybrid rocket motor. Good numerical predictions as compared to experimental performance data are presented. Computation of the N 2 O flow rate control with a ball valve shows smooth variation behavior which is desirable in propulsion system designs. NomenclatureC 1 ,C 2 ,C 3 ,C µ = turbulence modeling constants, 1.15, 1.9, 0.25, and 0.09. C p = heat capacity D = diffusivity H = total enthalpy K = thermal conductivity k = turbulent kinetic energy L/S = ratio of long-axis to short-axis Q = heat flux T = temperature t = time, s u = mean velocities V 2 = ∑ u 2 x = Cartesian coordinates or nondimensional distance α = species mass fraction ε = turbulent kinetic energy dissipation rate θ = energy dissipation contribution µ = viscosity µ t = turbulent eddy viscosity (=ρC µ k 2 /ε) Π = turbulent kinetic energy production ρ = density σ = turbulence modeling constants, 0.9, 0.9, 0.89, and 1.15 for Eqs.(2), (4), (5), (6). τ = shear stress ω = chemical species production rate

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Multiphysics simulations of rocket engine combustion

Cited by 44 publications

References 9 publications

N2O-HTPB Hybrid Rocket Combustion Modeling with Mixing Enhancement Designs

N2O-HTPB Hybrid Rocket Combustion Modeling with Mixing Enhancement Designs

Detailed kinetic computations and experiments for the choice of a fuel–oxidiser couple for hybrid propulsion

Multiphysics Computational Modeling of Hybrid Rocket Combustion

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