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 • 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.
Heat flux calculation Heat flux from the combustion products: Convective heat flux: Heat flux through the wall: Heat flux from the wall: Radiation heat flux: Heatbalance:
Numerical solution transient equation ofheat conductivity Two-dimensional transient equation of heat conductivity: - partial derivative of temperature depends from time; - thermal-conductivity coefficient; - density; - radius - heat capacity; - axis coordinate -declivity angle Difference analogue: Computational grid of combustion chamber
Difference analogue of boundary conditions Boundary conditions: Difference analogue: End nozzle: End nozzle: Surface of mixture head: Surface of mixture head: External surface of wall: External surface of wall: Internal surface of wall: Internal surface of wall:
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=0.11 m X=0.09 m
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).
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 • 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 • 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. • 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 • 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 temperature of combustion products ≈2800K for any propellant), 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.