The authors are developing a hybrid kick motor for small share-riding spacecraft. The authors launched the venture company Letara to commercialize hybrid kick motors in 2020. Critical technologies for practical use are fuel regression history prediction, nozzle erosion rate prediction, and ignition technology capable of re-ignition. This paper first describes why hybrid kick motors are necessary and how the authors started this development. The nitrous oxide the authors employed as a liquid oxidizer has a vapor pressure above the combustion chamber pressure. As a result, the combustion flow field changes according to the injection conditions, changing the axial distribution of the fuel regression velocity. The successful measurement of nozzle erosion rate history, which has been difficult in hybrid rockets, has allowed the authors to elucidate the mechanism of nozzle erosion. The introduction of a re-ignitable ignition system appropriate for small spacecraft with limited resources follows. Finally, the development status of a small-scale hybrid kick motor is presented.

This article presents the effects of swirl flow on combustion. It also discusses ongoing research on hybrid rocket engines using swirl flow. Swirling oxidizer injection proves to enhance fuel regression rate and combustion efficiency. Despite current efforts to clarify the flow field of swirl combustion, through experiments and CFD analysis, predicting the local fuel regression rate and its time variation along the fuel grain port remains an unresolved issue. Swirl flow has been applied to various hybrid rocket engines, including O/F shift suppression and thrust control by altering the swirl intensity during combustion, combination with low-melting-point temperature fuel for higher fuel regression rate, port shape design using additive manufacturing, and application of liquid oxidizer to the swirl combustion system. Several combustion issues still need to get solved.

In this study, combustion experiments were conducted to assess the feasibility and performance of a gas hybrid rocket system using Glycidyl Azide Polymer (GAP) as a gas generator fuel. Two types of gas generators with 60 mm and 80 mm diameter motors were investigated. Combustion tests were performed with end-burning grains, and the temperature inside the motor was measured using a φ1.0mm K-type thermocouple. The pressure range was from 3 to 10 MPa. The motor tests revealed that gas temperatures were approximately 80 K higher than those in strand tests. Both temperatures, however, were significantly lower than the adiabatic temperature. The efficiency of C* (ηC*) ranged from 0.7 to 0.85 depending on pressure and L*. A secondary combustion chamber was attached to an 80mm diameter motor, and a gas hybrid rocket was operated using gaseous oxygen. In this experiment, at a specific L/D ratio, the C* efficiency of the secondary combustion chamber reached 98 %. Based on the combustion experiment results of the gas hybrid rocket, it was demonstrated that the gas hybrid rocket system has a high C* efficiency. Furthermore, by optimizing gaseous oxygen (GOX) injection, there is potential for downsizing the secondary combustion chamber.

This paper introduces the analytical methods of thermal analysis for hybrid rocket fuel, and of evolved gas analysis for flash pyrolysis products. The thermal characterization of solid fuels requires the evaluation of phase changes, pyrolysis products, and thermophysical properties. This study performed measurements using conventional analytical chemistry methods and qualitative measurements of pyrolysis products. Furthermore, the temperature profile near the fuel surface was obtained because accurate measurement of the combustion flame temperature is essential for evaluating the propellant performance. The thermal, evolved gas, and kinetic analyses were conducted to investigate the phase-change process on the fuel surface. Because multicomponent polymeric materials, such as LT fuel, can lead to complex responses, this study evaluates the phase change behavior in detail by directly observing the color and shape changes of the heated samples. To measure the flash pyrolysis products generated in a flash heating environment, direct and real-time measurements are essential because the pyrolysis products generated from the burning surface are highly active. Qualitative analysis was carried out using a Curie-point pyrolyzer and an ion attachment-mass spectrometer (IA/MS) with a skimmer interface to reduce fragmentation and the second reaction of the products.

Hybrid rockets that employ non-explosive propellants are effective in enhancing the resilience of space transportation systems. For practical application of hybrid rockets, it is important to establish a safety evaluation method that considers their characteristics. Therefore, this paper summarizes the history of past and present research on safety evaluation methods for hybrid propellants. The most potentially hazardous situation to which hybrid propellants may be exposed is when the fuel is fragmented by some impact and a combustible mixture of leaked oxidizer and dust is formed. Therefore, the fragmentation of the fuel materials that can be employed in hybrid propellants is evaluated. Fuel fragmentation in a conditioned atmosphere is used to evaluate the occurrence of combustion. In addition, the combustible zone of the fuel dust is verified by the differences in dust concentration, particle size, and oxidizer concentration.

This article introduces and compares two types of dust-cloud combustion models: continuum and point-source models. The former model treats a dust cloud as if it were a continuum phase, while the latter considers heat generation from individual particles. Two dimensionless quantities are identified as the governing parameters in the limit of lean dust cloud of small particles: the dimensionless reaction time, τ_c, equivalent to the reciprocal of the Damköhler number, and the dimensionless ignition temperature, θ_ign. This paper focuses on predictions obtained for quasi-one-dimensional propagation. Without heat loss, extinction occurs when the adiabatic combustion temperature equals the ignition temperature; the dust-cloud concentration at extinction is independent of particle size. On the other hand, heat loss makes the extinction concentration depend on particle size. The continuum model is a good approximation of the point-source model when τ_c≳1, while its deviation from the point-source model is apparent when τ_c is smaller. The influence of particle-size distribution is discussed based on the point-source model.

Soot produced during combustion and carbon black used as an additive for rubber are carbon nanoparticles, which are aggregates of primary particles. Simulations that consider the shape changes of the aggregates are important not only for clarifying the determinants of soot and carbon black properties, but also for comparing the experimental particle size distribution with that obtained by simulation. In this study, we proposed a method to reduce the computational cost of a two-dimensional sectional model that can consider particle shape information and coupled it with a detailed kinetic mechanism. The importance of the coalescence rate model was also clarified by considering the coalescence rate model and comparing it with the particle size distribution obtained from experiments.

An experimental study was conducted to clarify the basic single-orifice pre-chamber jet combustion characteristics of a small two-stroke engine. In the experiments, the engine was operated at MBT with pre-chamber volume ratios of 3, 5, and 7 ％. The experimental equipment was a two-stroke, air-cooled, single-cylinder gasoline engine with a displacement of 63.3 cm³. The pre-chamber jet combustion flame was photographed through a quartz observation window, and in-cylinder pressure was measured at the same time. The results showed that the larger the volume ratio, the higher the intensity of the jet generated, resulting in more rapid combustion in the main chamber. It was also found that cooling losses occur when the jet impinges on the wall surface. The obtained combustion images were binarized to calculate the jet flow progression rate, which was found to have a strong correlation with the ignition delay period. The higher the jet flow progression rate, the longer the combustion period.