우주비행체 적용 단일액체추진제 하이드라진 추력기의 최적설계 및 성능시험평가

Abstract

Liquid-monopropellant thrusters have been developed and applied to the attitude and velocity control of space vehicles since the earlier stage of the global space era. Particularly, hydrazine (N2H4) thrusters in which hypergolic fuel is directly injected onto a catalyst bed have the advantages of relatively simple structure, low development cost, long/stable propellant storability, clean exhaust products, and so on. Therefore, the hydrazine thruster has a wide, up-to-date application as an engine for the attitude control system (ACS) in a lot of space vehicles such as space launch vehicles (SLV), Earth-orbiting satellites, and interplanetary spacecrafts. Technologies relevant to the space development is a typical measure to the competitiveness of the national science and technology, but there has been no systematic study for the design and development of medium-/high-thrust level of thrusters in the domestic region. Absence of the design/analysis technology is directly related to the absence of developmental applicability for the diversification of the engine models needed for upper-level systems. Moreover, rocket engine-related technologies are classified as a dual-use technology (DUT) in the space-advanced countries and are strictly controlled by the E/L (export license) from the component level to the technology documents and data. It is also thought that the core-technology transfer among the countries is impossible. For these reasons, a research has been conducted to secure the design-essential technology of liquid-monopropellant hydrazine thruster, and its ground-firing test results are presented in this dissertation with a part of development process. Development model (DM) was designed to produce 67 N (15 lbf) of nominal steady-state thrust at a propellant inlet pressure of 2.41 MPa (350 psia). For the design optimization and database (DB) construction of a hydrazine thruster, a flange-type assembly feature was implemented for varying the inner diameter and length of thrust chambers. Monopropellant-grade hydrazine was adopted as a propellant for the hot-fire test (HFT), and double-stage catalyst beds were filled with iridium-coated granular alumina (Ir/Al2O3) catalyst which is of 18-30 mesh size. An evaluation test for DM was performed on various thrust chambers having the characteristic lengths of 2.35, 2.42, 2.58, 2.65, 2.72, 2.79, 2.95, 3.13, and 3.58 meters, respectively. Performance verification of the DM engines were made through thrust, propellant flow rate, specific impulse, impulse bit, pressure and temperature in thrust chamber, hydrazine decomposition rate, thermal barrier efficiency of head-end assembly, and thrust response characteristics. First of all, the performance of DM thrusters reached about 95% in the characteristic velocity efficiency compared to their ideal requirements. Results of HFT indicate that an increment of the chamber characteristic length from the baselined model negatively affects the performance of the DM thruster: this seems to be primarily caused by excessive ammonia dissociation in the test conditions specified. The increase of chamber length also deteriorates propellant consumption efficiency with the supply pressure decrease. A supplement study was carried out to make a close examination into the pressure instability arisen from the DM. Low frequency instability (~ 50 Hz) was inherent in the test models, and in addition, increase of the chamber diameter and decrease of the operating pressure led to an amplification of pressure oscillation in the test conditions given. It was also verified through the HFT that the pulse response performance of DM showed superior characteristics to the commercialized hydrazine thrusters and its performance repeatability (pulse to pulse) was excellent, too. Finally, the oscillation hampered the pulse response performance of test models. Through this development study, the core technology with regard to the hydrazine-thruster design/development with its performance evaluation technique has been secured. The applicability of this research outcomes are so definite and the contributions are expected to the following areas: ● localization of the 3-axis attitude-control thruster system which might be equipped on the 1st, 2nd, and 3rd stage of SLV series ● development of apogee kick engine (AKE) system in GEO (geosynchronous Earth orbit) satellites ● securing a development technology of the propulsion system for interplanetary spacecraft, moon orbiter, landing module, and reentry vehicle, etc. ● a self-sustained growth of space technology in Korea ● development of the military technology and/or dual-use technology