Propulsion technologies

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Recent trends in the space community for smaller, cheaper and more frequent space missions have driven the development of micro- and nanosatellites. The use of small spacecraft constellations is an attempt to enhance the overall performance of communication and remote sensing tasks currently done by a relatively small number of large platforms. Because micro-technologies have the advantage of reducing the total resources required on a spacecraft, the continued development of micro-technologies for space applications will further enable small satellite missions. Nanosatellites (mass between 1 and 10 kg) impose significant limitations on mass, power and volume available for all subsystems including propulsion. In recent years, micropropulsion systems have been developed to address this need. A wide array of concepts will require the expansion of propellant gases through microscale nozzles. Because many micropropulsion systems will also operate at relatively low pressures, the investigation of low-Reynolds-number nozzle and jet flows has become increasingly important for realistic evaluation of new concepts.

When in orbit, spacecraft require on-board or secondary propulsion systems to perform orbit transfer, orbit maintenance, and attitude control maneuvers. An important issue in the use of any spacecraft propulsion system involves the assessment and reduction of effects caused by the interaction between the thruster plume and spacecraft surfaces. Direct impingement of a thruster plume on surfaces can generate unwanted torques, localized surface heating, and surface contamination. Self-impingement (i.e. the impingement of a thruster plume on a host satellite surface) generally occurs for small surface angles with respect to the propulsion system’s thrust vector or occurs in the thruster back flow. Cross impingement (i.e. the impingement of one spacecraft’s thruster plume onto another spacecraft) can occur at essentially any angle and is becoming increasingly important with the advent of microsatellite constellations. Although there have been few reported incidents of spacecraft failing because of spacecraft interaction effects from the secondary propulsion system, the tighter margins of contemporary spacecraft operation require improved optimization of all spacecraft systems. Improved assessment of the interaction potential of a thruster may also allow propulsion systems on spacecraft that are extremely sensitive to contamination concerns. Accurate prediction of plume interaction effects is technically challenging. It is very difficult and expensive to recreate the in-orbit environment within a ground-based laboratory facility. In terms of modeling, the flow physics associated with chemical and plasma propulsion systems is complex, requiring the development of advanced numerical techniques. In the past, due to the lack of detailed information, the prevention of the plume effects described here has been achieved primarily using very conservative spacecraft design. In the past few years, significant advances have been made in terms of both experimental facilities and numerical methods for the measurement and computation of the plumes of spacecraft secondary propulsion systems.

Therefore, this is an appropriate time to consider advancing the current state-of-the-art in numerical modeling for this important area of spacecraft design. This proposal seeks to address spacecraft-thruster interaction issues for both chemical and electric propulsion systems. As new thruster technologies are developed, investigations of spacecraft-thruster interactions will continue, and as new spacecraft technologies are developed, these studies will become increasingly important. The present project will involve numerical studies of exhaustion of gas jets of real spacecraft thrusters into vacuum. Such studies are necessary to estimate contamination and the action of these jets both on the spacecraft itself and on another spacecraft during orbital docking and undocking. Jet exhaustion occurs in a rapidly changing sequence of flow regimes: from continuum (inside the nozzle) through transitional (near the nozzle exit) to free-molecular (at a large distance from the nozzle exit). The flow is usually in an essentially non-equilibrium state. Numerical investigations will be performed by using two approaches: solving the Navier-Stokes equations inside the nozzle and using the DSMC method outside the nozzle. As the jet is sufficiently dense near the nozzle exit, the DSMC requires significant computational resources (memory and speed). Therefore, it is necessary to use massive parallel computers and a multizone approach where the results of computations in a certain domain around the nozzle are used as the boundary conditions for computations in a greater domain. The flow generated by interaction of several jets will be simulated. Efforts will be put into development of an approximate analytical model which will allow for determination of plume far field parameters using nozzle exit data and to assess the contamination of the spacecraft/satellite surface.

An alternative concept that can be used to fulfill on-orbit propulsion tasks (such as orientation and orbit correction) are electric thrusters. Their basic advantage over thermochemical control engines is a higher jet exhaustion velocity and, correspondingly, a lower flow rate of the working gas at an identical thrust (i.e., a higher specific impulse). One of the major problems of using these engines is the interaction of the plasma jet from the engine with the surfaces of solar arrays, radiators, and other sensitive elements of the space vehicle. A typical feature of the jets of such engines is an extremely high degree of thermal non-equilibrium of the flow, which necessitates the use of the kinetic approach based on the particle-in-cell/DSMC technique. Problems of exhaustion of jets from such engines into the space and interaction of these jets with the vehicle surface will be considered within the framework of this project. In particular, the PIC module for treating the charged particles will be implemented into the proposed DSMC GPU code. A state-of-the-art hydrodynamic model will be employed to simulate electrons while ions will be treated as particles in the conventional particle-in-cell/DSMC fashion. Modeling of the electric thruster plumes and their interaction with a 3D satellite/spacecraft models will be performed in order to assess possible impairment of communication systems, scientific instrumentations and solar arrays. A computational tool for prediction of electric thruster plume/satellite interaction will be developed as a deliverable in the framework of the proposed project. Special attention will be paid to validation of the physical and chemical models of the electric thruster plumes against ground-based experiments.

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