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With the aim of enhance the designability of burning surface, star and wagon wheel fuel types are applied extensively in hybrid rocket motors. However, mechanism of nozzle ablation in hybrid rocket motors with these complex fuel types is unclear, and this increases the difficulty of predicting ablation laws and thrust. This paper intends to investigate spatial distribution characteristics of nozzle ablation in the hybrid rocket motor with various fuel types. According to the principle that the grain port area and grain length are equal, tube, star, single-port wagon wheel, and multi-port wagon wheel grains are designed. Numerical models for combustion flow and thermochemical erosion are established, and three-dimensional numerical simulations are conducted. A firing test for nozzle erosion is conducted, and hydroxyl-terminated polybutadiene and 95 % hydrogen peroxide are employed. Errors of throat erosion rate and combustion chamber pressure among simulation results and experimental data are 4 % and 4.25 %, respectively. Calculation results demonstrate that the circumferential distributions of nozzle wall temperature and erosion rate show an obvious correlation with fuel types. The maximum throat erosion rates for star, single-port wagon wheel, and multi-port wagon wheel grains occur at the circumferential position of 0° (star slot, spoke slot, and midpoint of outer edge of wheel hole), while the minimum throat erosion rates locate at the circumferential position of 36° (star root, spoke root, and midline of adjacent wheel holes). The circumferential non-uniformity of throat erosion rate of multi-port wagon wheel grain is the highest. When the grain port area and oxidizer mass flow rate are the same, the throat erosion rate of multi-port wagon wheel grain is the highest, followed by the star grain, and the throat erosion rate of single-port wagon wheel grain is the lowest. • Effect of complex fuel types on spatial characteristics of nozzle erosion is studied. • Firing test of carbon-based nozzle erosion is performed to verify numerical models. • Three-dimensional simulations on erosion in hybrid rocket motors are conducted. • Simulation results agree well with test data. • Ablation mechanism of star, single-port, and multi-port wagon wheel grain is revealed. [ABSTRACT FROM AUTHOR]

The hybrid rocket motor (HRM) emerges as a promising propulsion system offering thrust adjustment at a low cost. Prior research has indicated that the inclusion of micro Al additives enhances HRM performance, yet a comprehensive investigation remains absent. Thus, this study aims to explore the impact of Al particle additive content and size variation on the characteristics of hybrid rocket motors (HRMs) through experiments and simulations. Firing tests are conducted on HRMs employing unadulterated HTPB, 38 % additive Al (2 μm), and 58 % Al (2 μm). Experimental findings reveal enhancements of 12.6 % and 33.9 % in regression rate, and 2.5 % and 8.0 % in specific impulse for the 38 % Al and 58 % Al fuel configurations, respectively, compared to the base case. Subsequently, a numerical two-phase combustion model is built to analyze the inner flow field characteristics. The high consistency between experimental and simulation results validates the efficacy of the numerical approach employed. Following this, a series of simulations covering Al content ranging from 0 % to 58 % and Al particle sizes from 0.4 μm to 75 μm is conducted. Simulation outcomes indicate that specific impulse and combustion efficiency enhancements correlate with the increase in Al mass fraction for particles with diameters under 35 μm. However, high Al content proves detrimental to specific impulse and combustion efficiency when particle diameter exceeds 35 μm. Additionally, regression rate and thrust demonstrate an increase with Al content, albeit the rate of increment diminishes with larger Al particles. This study comprehensively examines the influence of Al particle addition on HRM performance and elucidates the underlying mechanisms, thereby offering insights into HRM design and manufacture. • Hybrid rocket firing tests are performed on HTPB, 38 %, and 58 % Al particle fuel. • Simulations with various Al particle contents and diameters are performed. • High Al content with diameter under 35 μm is beneficial for combustion efficiency. • High Al content with diameter over 35 μm is detrimental for combustion efficiency. • Large particle reduces the thrust enhancement brought by Al content rise. [ABSTRACT FROM AUTHOR]

Precise regulation of oxidizer mass flow rate is a crucial technology for hybrid rocket motors, and a flow control valve serves as a core component of an oxidizer delivery system. In this paper, we designed a variable area cavitating venturi (VACV) as a flow control valve to control oxidizer mass flow rate. Three types of experiments were carried out to investigate VACA-based flow feedback control and its application in hybrid rocket motors. Specifically, we measured flow characteristics at different pintle strokes and conducted five cold tests and one ground hot test based on flow feedback control. Test results are as follows: first, the VACA flow characteristic curve shows a wide regulation range with a flow regulation ratio of 21.8 and performs good linearity of 96.76%. Second, the mass flow rate curve measured with flow feedback control tracks the target curve. Flow feedback control effectively improves control accuracy by maintaining the steady-state error within 1.2%. Third, hot test results demonstrate that the VACA with flow feedback control can effectively maintain the accurate and consistent oxidizer mass flow rate during the firing stage of the hybrid rocket motor. The findings of this research indicate that VACA with flow feedback control can be used as an effective flow regulation component to maintain a constant and accurate oxidizer mass flow rate during the working stage of a hybrid rocket motor. • A firing test of a hybrid rocket motor based on feedback control was conducted. • VACV shows good linearity with a wide regulation range in flow characteristics curve. • Feedback control with VACV improves accuracy, keeping steady-state error within 1%. • Feedback control maintains an accurate and consistent oxidizer mass flow rate. [ABSTRACT FROM AUTHOR]

This paper attempts to research ignition delay time and ignition process of catalytic hybrid rocket motor (HRM) with 95% hydrogen peroxide (95HP)/HTPB-based aluminum additive fuel. Firing tests are conducted with 95HP and different fuels on lab-scale HRMs. The aluminum weight fractions of four HTPB-based fuels are 38%, 48%, 58% and 65% respectively. Corresponding two-dimensional axisymmetric transient simulations are established to verify the relationship of catalytic ignition delay to fuel grain aluminum weight fraction and to catalytic decomposition temperature. The mono-propellant catalysis experiments of 95HP are carried out and show that the maximum temperature of decomposition products can reach 1140 K. The catalytic ignition delay time of 95HP/HTPB-based fuels is affected by the aluminum additive. When the aluminum weight fraction of the fuels is increased from 38% to 65%, the effective heating time of the ignition is correspondingly increased from 0.250 s to 0.556 s. The flame pictures are taken by the high-speed camera in the firing tests. A large number of sparks appear in the flame when it reaches a stable state, but there is no spark in the flame at the initial stage of ignition. Therefore, it is indicated that the aluminum does not participate in the chemical reaction at the initial stage of ignition, but may change the physical properties of the fuel. Similarly, the simulation results prove that the increase of the aluminum weight fraction of the fuel will increase the ignition delay time while the increase of catalytic decomposition temperature will greatly reduce the ignition delay time. • Firing tests are conducted with HTPB-based aluminum additive fuel. • Increasing of the aluminum fraction of fuel will increase the ignition delay time. • Aluminum does not participate in the chemical reaction before the fuel is ignited. • Two-dimensional axisymmetric transient simulations are conducted. [ABSTRACT FROM AUTHOR]

This paper aims at the dynamic characteristics study of variable thrust HRMs, for which a two-dimensional transient numerical model of HRM is established. Complete transient simulation based on dynamic mesh has been performed sequentially. An HRM with a thrust of three stages is designed and tested. The pressure, regression rate and thrust of the simulation results are in good agreement with the experiment data, which verifies the feasibility of the dynamic mesh technique for fuel regression simulation in HRMs. Specifically, the maximum deviations of pressure and thrust are less than 3.6% and 3.2%, respectively at the large mass flow rate stage. The numerical simulations show that the regression rate is not only related to the oxidizer mass flux, but also affected by the dynamic regression process. Max deviation of regression rate is 35% at the same oxidizer mass flux. Moreover, it is also demonstrated that the presence of vortex near front edge of fuel grain has a great effect on motor performance. Regression rate at front part of solid fuel is significantly high at the initial stage of throttling process, and regression rate downstream of the vortex obviously increases with time. In addition, the gas temperature of the pre-combustion chamber continuously increases with regression of fuel surface. Gas temperature of the post-combustion chamber is positively correlated with the oxidizer mass flow rate. This research can predict performance of HRMs more accurately, and provide a valuable guidance for variable thrust HRMs design. • A numerical simulation model of variable thrust hybrid rocket motor is established. • Dynamic mesh technique is used to simulate the regression of solid fuel surface. • A firing test of variable thrust hybrid rocket motor are conducted. • Average regression rate is different at the same oxidizer mass flux. • Regression rate of downstream of the vortex obviously increases with time. [ABSTRACT FROM AUTHOR]

Hybrid rocket motors (HRMs) easily achieve variable thrust by changing oxidizer mass flow rate. This paper represents numerical and experimental studies on the throttleable HRM with aerospike nozzle, which can enhance the performances of the throttleable HRM theoretically. By comparing the performance of the throttleable HRM with the aerospike nozzle and that with the de Laval nozzle, the characteristics of the aerospike nozzle applied to the throttleable HRM are explored. Both numerical and experimental investigations are conducted. In the simulation part, characteristic velocity, combustion efficiency, thrust coefficient and specific impulse of the HRM with the two types of nozzles at different oxidizer mass flow rate are obtained. Comparing with the HRM with the de Laval nozzle, the HRM with the aerospike nozzle can improve specific impulse by about 8.9%∼23.4% when the working pressure ratio is lower than the design pressure ratio. The combustion efficiency of the throttleable HRM with the aerospike nozzle is 3.9%∼8.6% higher than that with the de Laval nozzle because of the core structure of the aerospike nozzle. When the oxidizer mass flow rate is 50 g/s, the thrust coefficient of the aerospike nozzle is increased by 12.3% relative to that of the de Laval nozzle which flow separation occurs in. The corresponding firing tests are performed by the lab-scale motor with 90% hydrogen peroxide (HP) and polyethylene (PE) grain. The numerical and experimental results agree well. The combination of the throttleable HRM and the aerospike nozzle has the advantages to improve the performance of HRMs. [ABSTRACT FROM AUTHOR]