Browsing by Author "Wei, S. S."

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COMPARISON OF THE PROPULSION PERFORMANCE OF AEROSPIKE AND BELL-SHAPED NOZZLE USING HYDROGEN PEROXIDE MONOPROPELLANT UNDER SEA-LEVEL CONDITION
(2019-06-01)
Lai, A.
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Wei, S. S.
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Lai, C. H.
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Chen, J. L.
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Liao, Y. H.
This study investigates numerically the performance of applying aerospike nozzle in a hydrogen peroxide mono-propellant propulsion system. A set of governing equations, including continuity, momentum, energy and species conservation equations with extended k-epsilon turbulence equations, are solved using the finite-volume method. The hydrogen peroxide mono-propellant is assumed to be fully decomposed into water vapor and oxygen after flowing through a catalyst bed before entering the nozzle. The aerospike nozzle is expected to have high performance even in deep throttling cases due to its self-compensating characteristics in a wide range of ambient pressure environments. The results show that the thrust coefficient efficiency (C-f,C-eta) of this work exceeds 90% of the theoretical value with a nozzle pressure ratio (PR) in the range of 20 similar to 45. Many complex gas dynamics phenomena in the aerospike nozzle are found and explained in the paper. In addition, performance of the aerospike nozzle is compared with that of the bell-shape nozzle.
Experimental investigation of the effect of nozzle throat diameter on the performance of a hybrid rocket motor with swirling injection of high-concentration hydrogen peroxide
(2019-11-01)
Wei, S. S.
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Lee, M. C.
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Chien, Y. H.
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Chou, T. H.
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Wu, J. S.
One of the issues of using a normal graphite nozzle for the hybrid rocket propulsion is the serious throat erosion due to combustion with a higher O/F ratio. This may undermine the throttling capability of the hybrid rocket engine. In this study, we would like to address how the propulsion performance changes under the conditions of different nozzle throat diameters and O/F ratios. We designed and tested a 40-kgf class single-port hybrid rocket motor with the swirling injection of oxidizer, which utilizes the 90 wt% hydrogen peroxide and polypropylene (PP) as oxidizer and fuel, respectively. Three different diameters of graphite nozzle throat (10, 11, and 12 mm) were used to approximate various conditions of nozzle erosion, while the mass flow rates of the injected oxidizer were kept the same. The hot-fire test results indicated that the thrust was nearly the same even the nozzle throat diameters and O/F ratios were different. Especially, the discrepancies among the measured thrusts and oxidizer ISP under various test conditions were found to be less than 1%, which is highly beneficial for the purpose of the thrust control using mass flow rate control of the oxidizer for a hybrid rocket engine with swirling injection of high-concentration hydrogen peroxide.
NUMERICAL INVESTIGATION OF A N2O/HTPB HYBRID ROCKET MOTOR WITH A DUAL-VORTICAL-FLOW (DVF) DESIGN
(2017-12-01)
Lai, A.
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Lin, Y. C.
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Wei, S. S.
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Chou, T. H.
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Lin, J. W.
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.