A compact hybrid rocket motor design that incorporates a dual-vortical-flow (DVF) concept is proposed. The oxidizer (nitrous oxide, N2O) is injected circumferentially into various sections of the rocket motor, which are sectored by several solid fuel “rings” (made of hydroxyl-terminated polybutadiene, HTPB) that are installed along the central axis of the motor. The proposed configuration not only increases the residence time of the oxidizer flow, it also implies an inherent “roll control” capability of the motor. Based on a DVF motor geometry with a designed thrust level of 11.6 kN, the characteristics of the turbulent reacting flow within the motor and its rocket performance were analyzed with a comprehensive numerical model that implements both real-fluid properties and finite-rate chemistry. Data indicate that the vacuum specific impulse (Isp) of the DVF motor could reach 278 s. The result from a preliminary ground test of a lab-scale DVF hybrid rocket motor (with a designed thrust level of 3,000 N) also shows promising performance. The proposed DVF concept is expected to partly resolve the issue of scalability, which remains challenging for hybrid rocket motors development.

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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One main physical feature of hybrid rocket combustion is its diffusion flame nature that has required excessively long combustion chamber which 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 fuel regression rate and combustion efficiency. Thus, shortened combustion chambers can be used to deliver reasonable thrust performance of hybrid rockets. In the present study, a compact hybrid rocket motor design is investigated to provide a form factor with small slenderness, which is suitable for improving the overall system designs for hybrid rockets. This design concept features in multiple vortical flow structures such that much enhanced combustion efficiency can be obtained. A 3-D computational model with finite-rate chemistry using reduced kinetics mechanism and radiative heat transfer effects are employed in the present investigation to assess the mixing effectiveness and combustion efficiency of the present design. This computational model has been validated for a wide range of rocket propulsion design problems, including single-port hybrid rocket motors with mixing enhancement vortex generator devices. The internal ballistics and flame structure in the present multiple vortical-flow hybrid rocket engine designs using HTPB fuel with nitrous oxide or hydrogen peroxide oxidizers are investigated. NomenclatureC 1 ,C 2 ,C 3 ,C  = turbulence modeling constants, 1.15, 1.9, 0.25, and 0.09.

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