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A feasibility study related to the scale-up of solid fuel with multiport was performed using the regression rate equation derived from a laboratory-scale hydrogen peroxide (H2O2) hybrid rocket usin...

This paper presents a thin broadband microwave absorber that is manufactured by electrical modification of a glass fabric using two different types of coating techniques in the C band to the Ku band (4–18 GHz). Unlike the production method of conventional absorbers, the proposed microwave absorbers were manufactured using a triple-layered structure. The latter consists of glass/epoxy, Ni-coated glass fabric, and Ni-Co-coated glass fabric without dispersed carbonaceous nano-conductive particles in the polymer matrix. The glass fabric was coated with a Ni and a Ni-Co magnetic nanomaterial via electroless plating and RF magnetron sputtering techniques. The element components, phase identification, and magnetic properties of the conductive and magnetic material-coated glass fabrics were examined by energy-dispersive X-ray spectroscopy, X-ray diffraction, and vibrating sample magnetometer, respectively. The results showed that high dielectric and magnetic properties can be obtained in the target frequency range owing to the conductive and the magnetic material. The proposed two types of broadband microwave absorbers have a low thickness value of 0.858 and 3.305 mm; the design of the broadband absorbers resulted in two strong resonance peaks near 8.5 and 16.3 GHz and satisfied the −10 dB absorption performance from 6.32 to 18 GHz.

This paper reports a feasibility study of regenerative micro-cooling channels in a liquid microthruster composed of thermally fragile materials. Glass, which is among the most thermally insulating materials, has been used as microthruster fabrication material to suppress excessive heat loss in micro scale thruster. However, the frangibility of glass has remained a challenge to be solved. To thermally manage the fragile structure, the use of regenerative micro-cooling channels in a microthruster is suggested in this work, and the feasibility was tested through design, fabrication and experimental performance of a glass microthruster with microchannels. Nine photosensitive glass layers were wet etched and integrated to fabricate the microthruster. Before integration of the layers, a fabricated Pt/Al 2 O 3 catalyst was inserted into the chamber of the microthruster for propellant decomposition. Hydrogen peroxide (90 wt%) was used as a monopropellant and served as the working fluid for regenerative cooling. A liquid microthruster with micro-cooling channels was successfully fabricated with a photosensitive glass MEMS process. Experimental performance tests were conducted while measuring the microthruster chamber pressure, chamber temperature, and surface temperatures. The test results showed normal operation of the microthruster, which had an estimated thrust of approximately 48 mN and temperature efficiency of approximately 41%. The decreasing surface temperatures of the microthruster during thruster operation successfully validated the cooling effect of the micro-cooling channels and demonstrated their practicality for the regenerative cooling of liquid microthrusters.

A novel experimental method to evaluate the ignition characteristics of hypergolic propellants was investigated. Three hypergolic liquid fuels with different ignition characteristics were prepared, and their reaction rates with hydrogen peroxide were measured in a microreactor fabricated using a microelectromechanical systems lithography process. To prevent evaporation and boiling by the reaction heat, a low-concentration of hydrogen peroxide of 34.5 wt % was used as a reference oxidizer. The reaction rates of the hypergolic fuels were obtained from temperature measurements in the microreactor and correlated with the ignition delays that were separately measured by drop tests with high-test peroxide. The reaction rate measurements were fully compatible with the drop test results and proved to be useful as an alternative method for the selection of hypergolic propellants by ignition characteristics.

A feasibility study of the design of multiport solid fuel using the ratio of the nozzle throat area to the port area was performed in a hydrogen peroxide/high-density polyethylene hybrid rocket. The port diameters of the multiport solid fuel according to the change of the number of ports were designed by setting the ratio of the nozzle throat area to the port area at 0.25, 0.3, 0.5, and 1. To obtain the design point of the multiport solid fuel, the final port diameter and performance parameters were analyzed using a numerical estimation code based on the regression rate equation. The numerical estimation indicated that the solid fuel with nine circular ports with a length of 400 mm, and a port diameter designed with ratio of 0.5 between the nozzle throat area and the port area, had an oxidizer-to-fuel ratio (O/F) close to the design value and high specific impulse efficiency. To verify the value calculated using the numerical estimation, a 2500-N-class hydrogen peroxide/high-density polyethylene hybrid rocket with multiport solid fuel was designed, and experiments were performed. The specific impulse and overall regression rate calculated from the experiments and numerical estimation were compared. The maximum differences of less than 10.7% and 4.38% were observed in the comparison. Based on the small difference, the estimation results calculated using numerical estimation at ratios (the nozzle throat area to the port area) of 0.25, 0.3, 0.5, and 1 were demonstrated. The design method of the multiport solid fuel in a hydrogen peroxide/high-density polyethylene hybrid rocket using the ratio of the nozzle throat area to the port area and the numerical estimation based on the regression rate equation was feasible.

To develop sounding rockets that can be used for various missions, a university, research center, and two enterprises collaborated to conduct a critical-system development project on a small-scale sounding rocket. The Korea Advanced Institute of Science and Technology (KAIST) Rocket Laboratory and Space Solution Co., Ltd. collaborated to develop a hybrid rocket and a propellant-feeding system using hydrogen peroxide as an oxidizer. The KAIST Satellite Technology Research Center has developed a non-pyrotechnic recovery system, and Naraspace Technology Co., Ltd. has developed rocket avionics. Using a 2.39-m technology demonstrator that weighs 16.4 kg, the developed system has been validated at low altitudes, and the applicability of the small-scale sounding rocket has been evaluated. The combustion stability, short ignition delay, and high performance of the H2O2 hybrid rocket were confirmed through static firing and flight tests. The reliability and capacity of the non-pyrotechnic recovery system (which uses a spring) were confirmed by ground and flight tests. Perfect operation of the developed rocket avionics, such as control of the main oxidizer valve and recovery system, emergency functions, data transmission, and data storage was confirmed through ground and flight tests.

In regard to propulsion system applications, the stability of liquid propellants in long-term storage is of increasing importance, and this had led to a greater interest in gelation technology. As part of a preliminary test to determine the feasibility of using a gel propellant in a rocket with a catalyst bed, a hybrid rocket with a catalyst reactor using a gel propellant as an oxidizer was tested for the first time in this study. Experiments were conducted with two different oxidizers: one with liquid phase hydrogen peroxide and the other with gel phase hydrogen peroxide, as well as high-density polyethylene as fuel for a 250 N class hybrid thruster performance test. The thruster was designed with the catalyst ignition system, and a catalyst was manufactured to be inserted into the catalyst reactor to facilitate oxidizer decomposition. While the test result with neat hydrogen peroxide indicated sufficient decomposition efficiency using a manganese dioxide/alumina catalyst and successful autoignition of the fuel via the decomposed product, gel hydrogen peroxide exhibited insufficient decomposition and there were difficulties in operating the thruster as a part of the catalyst was covered in the gelling agent. This preliminary study identifies the potential challenges of using a gel phase oxidizer in a catalyst ignited hybrid thruster and discusses the technical issues that should be addressed in regard to a gel propellant hybrid thruster design with a catalyst reactor.

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A liquid monopropellant MEMS thruster was designed, fabricated and tested. For application on a nanosatellite for orbit control, a liquid propellant MEMS thruster delivers better performance than a solid thruster. Two issues must be addressed for a liquid monopropellant MEMS thruster: high energy content of the monopropellant to overcome the excessive heat loss associated with the small scale of the thruster, and repeatability of generated thrust force. The present study proposed blending 90 wt% hydrogen peroxide with ethanol at an oxidizer to fuel ratio of 30 to enhance the energy content of the propellant. The thruster structure was constructed using glass layers that were individually patterned by wet etching. The decomposition catalyst was separately prepared by wet impregnation of the active material, Pt, on the gamma alumina pellets and inserted into the thrust chamber before the UV bonding process of the glass layers. The firing test of the assembled MEMS thruster was conducted and thrust was measured both with ethanol blended hydrogen peroxide and pure hydrogen peroxide as a reference monopropellant. The measured thrusts were approximately 30 mN for both 1.7 ml min−1 flow rate of blended hydrogen peroxide and 2.0 ml min−1 flow rate of pure hydrogen peroxide. The measured thrust for 1.7 ml min−1 pure hydrogen peroxide was approximately 24 mN. The measured thrust was 40% less than the design thrust for both monopropellants. The uncertainty of the thrust was less with blended monopropellant than with pure hydrogen peroxide.

MEMS thruster with blended propellant was fabricated and experimentally tested in order to improve specific impulse of micro scale monopropellant thruster and to improve stability of thrust to be better. 90 wt. % H2O2 blended with 25 O/F ratio ethanol was used as propellant of thruster and platinum on alumina support was used as catalyst for decomposition of propellant. Thruster was made by five layers of photosensitive glasses. Four layers were integrated by thermal bonding method and catalyst was directly inserted into chamber before UV bonding process for the last layer bonding. Results of experimental tests showed ethanol blended hydrogen peroxide had higher specific impulse than unblended hydrogen peroxide. Expected improvement of thrust stability due to the blended propellant was found only in the transient state of thrust. Also, unlike the thrust instability of vertical type thruster of previous research, improvement of thrust stability was found owe to horizontal type thruster pattern on glass, despite aspect ratio limitation of glass fabrication with wet etching process. During the experimental test, combustion phenomena of ethanol with decomposed hydrogen peroxide were observed through glass layer and it made fracture on structure of thruster.