ISSP 2007

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Information about ISSP 2007
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Published on January 22, 2008

Author: Teresa1

Source: authorstream.com

Main directions of development of thrusters for reactive control systems of upper stages and spacecrafts.:  Main directions of development of thrusters for reactive control systems of upper stages and spacecrafts. Alexander A.Kozlov, Irena A.Bazanova, Aleksey G.Vorobiev, Igor N.Borovik International Symposium on Space Propulsion (ISSP2007) Main directions of development of thrusters for reactive control systems of upper stages and spacecrafts:  Main directions of development of thrusters for reactive control systems of upper stages and spacecrafts Using ecologically clean propellants with high energy. Estimation of the propellant efficiency from the final velocity of space vehicle criteria. Using the computational model for the research and design of thrusters of altitude control system (ACS). Development and application of different types of ignition systems. Perspective of using energy of solar and chemical power. Ballistic efficiency of the different propellants:  Ballistic efficiency of the different propellants Thrusters at ecologically clean propellants:  Thrusters at ecologically clean propellants Thrusters at ecological clean propellants Mathematical model of the heat state:  Mathematical model of the heat state Combustion chamber parameters: Internal and external geometry Combustion chamber material properties (density, specific heat capacity, heat conductivity, radiation coefficient) Thermocouple location Characteristics of thruster: Thrust Combustion chamber pressure, nozzle exit pressure Propellant Pressure of propellant at the mixing head entry Mixing parameters: Propellants mass flow rate Mixture ratio Number of injectors, pattern of injectors Film cooling injectors, film cooling mass flow rate, film cooling propellant Result parameters are: Mixture ratio in the wall layer Turbulence factor in the wall layer Gas temperature in the wall layer Temperature of inner surface of wall, temperature of outer surface of wall Transient heat flow through the wall Integral specific impulse Boundary conditions:  Boundary conditions Heat flux from the combustion products: Radiation heat flux: Convective heat flux: Heat flux through the wall: Heat flux from the wall: Heat balance: Mathematical model of the heat condition:  Mathematical model of the heat condition The calculated and experimental values of temperature of external wall for the engine ESTMAI-200 (F=200N, p=0.9 MPa, MON-UDMH, O/F=1.85, E(pk/pa)=1000, 1-injector head). X=0.09 m X=0.11 m Energy characteristics and optimal parameters of thruster:  Energy characteristics and optimal parameters of thruster The calculated values of specific impulse depending on maximum temperature of external wall in critical section and for the engine ESTMAI-200 (F=200N, p=0.9 MPa, MON-UDMH, E(pk/pa)=1000, 1-injector head). The calculated values of specific impulse depending numbers of injectors and for the engine ESTMAI-500 (F=500N, p=1 MPa, pressure (exit nozzle)=0.001 MPa, Hydrogen Peroxide- Kerosene). Characteristics of ignition systems:  Characteristics of ignition systems Electric spark:  Electric spark Input voltage - 27±3 V Output voltage - 12 kV Mass of transformer - 200 g Power consumption - 40 W. Clearance - 1 mm. Electric spark on the mixing head of 200N thruster (O2(gas)+Ker) during experiment. (High-power arc inside mixing pre-chamber.) Glow plug ignition:  Glow plug ignition Experimental combustion chamber of 20N engine with propellants kerosene+O2 with one center-placed two-component injector. Glow plug M3 with iridium glowing coil for model airplanes was inserted directly in cylinder wall of combustion chamber. Power consumption of the igniter is 6 W. Time delay for glowing coil is about 1 second. Design of 20 N engine with glow plug ignition. Glow plug ignition on 20 N engine Catalyst based ignition:  Catalyst based ignition In the development of thrusters, its became very attractive to use catalyst-dissolved kerosene for propellants kerosene+GO2m kerosene+H2O2, which guarantied self-ignition of these ecologically clean propellants. Solid catalyst GNII ChTEOS was used for decomposition hydrogen peroxide at two mixing heads (1 –injector and 7-injectors) for 200N engine (H2O2+kerosene). Catalyst was prepared basically with solid composition KMnO4 with special treatment of the surface of granules. The first test was conducted successfully and proved the ignition of kerosene with the product of decomposition H2O2 (concentration 94%). 200 N mixing head with catalyst based ignition. Solar-heated engine device with solar battery and electro-heated accumulator.:  Solar-heated engine device with solar battery and electro-heated accumulator. Tank with hydrogen Electro-pump of hydrogen Compressor Volume with gaseous hydrogen Tank with oxygen Electro-pump of oxygen Chamber of afterburner Photo-electrical battery Electro-chemical accumulator Transformer-control equipment Electro-heated accumulators Conclusions:  Conclusions At present most of spaceships and upper stages use storable toxic propellants (N2H4, (CH3)2N2H2, CH3N2H3, A-50, N2O4). However, the tendency was formed to the use of thrusters with ecologically clean propellants: O2+ker, O2+H2, O2+CH4, H2O2+ker. Ballistic efficiency of ecological clean propellants, (if the gas temperature in combustion chamber reached ≈2800K), surpasses that of traditional propellants N2O4+UDMH. The mathematical model of thruster’s heat state is developed. This model lets to decrease the numbers of fire tests and, therefore, its cost. It may be used for the optimization of thruster’s parameters at the early stage of design work. Different types of ignition systems of thrusters (electric-spark, glow plug, catalytic, gas-dynamic) are tested and the recommendations for its application were given. Space two-mode engine, using chemical (H2+ O2) and solar energy (thermal accumulator for the heating of hydrogen), developed in Russia, has big perspective for space transportation systems.

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