Electrode structure is a significant factor influencing the ignition and combustion characteristics of electrically controlled solid propellant. This study designs three types of fixed-electrode structures: inclined plate, top side, and staggered. When the voltage exceeds 150 V, the propellant reacts and ignites in all electrode structures. In the inclined plate electrode structure, the combustion of the propellant is nonuniform; in the top-side electrode structure, the combustion is unstable. In the staggered electrode structure, the propellant combusts uniformly and stably, and increasing the operating voltage of the propellant can enhance the initial current peak during electrification, thereby accelerating the ignition process. However, when the voltage reaches above 350 V, the current rapidly decays to zero after electrification, preventing the propellant from sustaining combustion. When the spacing between electrodes in the staggered structure is 8 mm, only localized combustion occurs; at 4 mm spacing, high-temperature flames can easily lead to breakdown discharge under high voltage, resulting in unstable combustion. In the parallel-plate electrode structure, the power of the propellant is higher, leading to more intense combustion. However, this structure has a longer ignition delay time and is less capable of completely extinguishing after powering off.

Among the many challenges to integrating a rotating detonation combustor (RDC) into a gas turbine (GT), cooling is one of the most predominant due to the high heat loads generated by the combustion process. To design a cooling system for an RDC that allows for its sustainable operation, a quantification of the heat loads of an RDC operating at conditions representative of GT is necessary. The presence of a detonation wave interacting with the boundary layer and a small annulus width leads to higher heat transfer when compared to a conventional GT combustor. This paper numerically models the heat flux and heat transfer coefficient of an RDC that would be relevant for setting cooling requirements in practical systems. Integral boundary-layer methods are employed to build a tool that uses 2D distributions of integral quantities to obtain the heat flux. A global heat transfer model is built using these simulations as a reference. The model was used to show that the heat transfer in the RDC is enhanced by the presence of a blockage at the combustor outlet. The effects on the flowfield are associated with an increase in chamber pressure and detonation strength, both causing an increase in the heat transfer coefficient to the liner walls.

In this study, the axisymmetric nozzle for the SW300B micro-turbojet engine is redesigned as an axisymmetric ejector nozzle, rectangular nozzle, rectangular ejector nozzle, rectangular lobed nozzle, and rectangular lobed ejector nozzle. The flow rate, thrust, fuel consumption rate, and infrared radiation (IR) characteristics of the micro-turbojet engine with each of the above-mentioned six nozzle types are studied experimentally. The effect of the ejector nozzle type on the aerodynamic performance and IR characteristics of the engine is then analyzed. The findings of this study reveal that ejector nozzles enhance the total flow rate of the engine, increase thrust, and reduce thrust-specific fuel consumption (TSFC). Under test conditions, the pumping characteristics of the rectangular lobed ejector nozzle are the strongest. Using this nozzle type, the ejector secondary-to-primary flow ratio reaches 48.65%, the engine thrust rises by 13.64%, and the TSFC falls by 12.02%. Moreover, the rectangular ejector nozzle effectively balances the thrust performance and IR characteristics of the engine. Compared to the turbojet engine with the rectangular lobed ejector nozzle, the thrust of the engine with the rectangular ejector nozzle is 4.7% higher. Additionally, the IR signature of the engine with the rectangular ejector nozzle is significantly lower than that with the axisymmetric ejector nozzle. In the 0° and 30° directions on the vertical detection plane, the IR characteristics fall by 20.2% and 47.7%, respectively.

Hybrid rockets offer various advantages, such as safety, controllability, and eco-friendliness, compared to solid and liquid rocket propulsion. Despite these, it has drawbacks such as low regression rate, lower combustion efficiency, and change in oxidizer-to-fuel ratio with burn time. Various methods have been adopted to overcome these issues. The present study analyzes the impact of adding a secondary swirl injector, along with the conventional head-end swirl injector, on the performance of a hybrid rocket motor. Paraffin wax was used as fuel and gaseous oxygen as the oxidizer. The secondary injector was positioned at three locations: 0.25, 0.50, and 0.75 secondary swirl location/entire motor length (X/L). Results showed that the secondary swirl injector prevented the swirling flow from decaying within the combustion chamber. For the head-end swirl injection, the tangential component of the swirl decays progressively after 200 mm along the flow direction. The optimal position for the secondary injector was found at 0.50 X/L, where the regression rate increased by 33.75%, combustion efficiency improved by 13.4%, and thrust and pressure increased by 53% and 30%, respectively, compared to the head-end swirl case. It was also observed that longer chambers led to incomplete combustion due to swirl decay, while shorter chambers gave a reduction in residence time, limiting the combustion efficiency.

Wind tunnel experiments were conducted on the concave and rectangular channels under incoming Mach numbers of 1.85 and 2.70. In the concave channel, the shock train structure leaned toward the concave wall, whereas a large flow separation existed on the flat wall. Sweeping the shock train leading edge abruptly increased the pressure, with the downstream pressure in the swept area increasing approximately linearly until it stabilized. The energy of the pressure fluctuations caused by the interaction between the shock train and boundary layer in the channel was concentrated in the frequency range of 2–1000 Hz. Compared to the rectangular channel, the shock train structure in the concave channel leaned toward the concave side with enhanced asymmetry. The anti-backpressure ability of the concave channel was stronger than that of the rectangular channel. The same change in the blockage ratio caused the shock train in the concave channel to move over a longer distance than in the rectangular channel. The stronger resistance to backpressure of the boundary-layer fluid in the concave channel results in stronger single shock waves in the shock train flowfield, with flow separation occurring closer to the downstream.

Shrouded stator cavity flow is a major source of aerodynamic loss that limits the compressor pressure rise capability. This paper analyzes the loss mechanisms of cavity leakage flow by examining the mean flow phenomenon and the entropy generation under a broad range of compressor operating conditions. In the first place, the effectiveness of the numerical method is validated using the experimental data of the detailed flowfield. Then, the disfavored mean flow phenomenon, the entropy-based loss of efficiency, and the irreversible entropy generation rate with and without the cavity are analyzed. Results show that the cavity generally reduces the compressor isentropic efficiency by about 0.5–0.9% in the investigated case. The extra entropy loss generation mainly comes from the stator passage rather than the cavity passage. The nature of the extra irreversible entropy generation is mostly turbulent viscous dissipation rather than the mean flow terms and the irreversible heat flux terms. The mean flow phenomena associated with the extra loss are mainly the partitioning of the stator inlet profile, the circumferential and radial migration of leakage flow, and the changes in the separated flow topology. These understandings provide useful guidance for the aerodynamic design of the shrouded stator.

The paper investigates two-dimensional slit flame front dynamics subjected to transverse and axial perturbations simultaneously, addressing the practical importance of azimuthal and longitudinal thermoacoustic oscillations occurring in annular combustion chambers. Specifically, responses of time-averaged flame fronts to two-way perturbations are theoretically investigated and analyzed using different disturbance propagation velocity models. Analytical solutions of time-averaged flame fronts with two-way flow disturbances are derived using the G-equation model and perturbation method, which incorporate undisturbed steady properties and disturbed zero-frequency properties. Time-averaged properties of flame configurations cannot be assumed as undisturbed steady properties, acknowledging the nonlinear nature of flame dynamics. The physical mechanisms of competing and coupling behaviors between transverse disturbance and axial disturbance have been investigated. Results highlight the substantial impact of the transverse-to-axial disturbance ratio on time-averaged flame fronts and heat release rates. Time-averaged flame heat release rate can be enhanced as the slit flame becomes longer, indicating that flame aspect ratio plays a pivotal role in shaping time-averaged flame front dynamics. Moreover, the disturbance propagation velocity model becomes a critical factor affecting time-averaged flame heat release rates, which will be enhanced with increasing transverse mean flow velocity.

In solid propellant rocket motors, the design of grain port geometry primarily dictates the motor ballistic performance. The flexibility offered by the spatial variation of the port shape on motor performance is often discounted due to the requirement of complex mandrel shapes along with their associated manufacturing issues coupled with propellant processing constraints. Literature data indicate improvement in motor volumetric loading and stability by planned distribution of burn surface area along motor length. This paper deals with the realization and testing of a subscale motor with central finocyl-based grain geometry. A novel method of casting complex grain shapes in solid rocket motors using the wax mandrel technique is discussed. The wax mandrel preparation, propellant casting procedure, and wax mandrel removal setup are elaborated. Static testing results of the subscale motor with central finocyl port geometry are also presented as proof-of-concept. This work has demonstrated the feasibility of realizing solid rocket motor grains using the wax mandrel technique, which further enables realization of complex grain designs necessary to meet various mission requirements.