## 4 Implementation of Spacecraft Using Interplanetary Medium
In the first approximation, the process of plasma capture and acceleration can be considered a continuous process for \(D\geq~{}200\) m when the space medium might serve as a matter tank for the plasma thruster. In this regard, two questions immediately arise: Is there enough radiation energy to create thrust? For how long shall the spacecraft accelerate? As already mentioned, the energy required for the acceleration and ionization of trapped particles is obtained from a solar battery situated on the surface of the device (as depicted in Figure 1). The final plasma flow velocity is determined by the radiation flux and the method adopted for plasma acceleration. For example, one can use the collective acceleration method based on the momentum transfer of the external electromagnetic field to the particles of the flow [35; 36; 37; 38] or some type of Hall thruster (see [3; 4; 5; 6; 7; 39; 40; 41; 42]). However, as our scheme necessitates various particle types of different masses, it is imperative to take into account the effects of mass and charge separation in crossed electric and magnetic fields (see [43; 44; 45]). In the present work, we do not focus on these points in detail, limiting ourselves to considering only the energy and material aspects of the discussed scheme (we hope to touch on these questions in future works). Nevertheless, as an illustration, we now consider one of possible technical approaches to the acceleration of the captured matter sketched in Figure 5[35; 36; 37; 38]. As is seen from Figure 5a, the flow of collected particles is injected into the acceleration section from the discharge chamber (if the captured matter is neutral particles). Figure 5b explains the physical idea of this plasma accelerator for a flow containing electrons, single-positive-charged ions, and heavy multiply charged negative particles. In this configuration, an alternating axial magnetic field, \(B_{z0}(t)\), generates an azimuthal component of the electric field \(E_{\varphi_{0}}(t)\) which, in turn, rotates the electron current, \(\dot{j}_{e\varphi}\), the dust particle current, \(\dot{j}_{d\varphi}\), and the ion current, \(\dot{j}_{ip}\), in different directions. Herein, the acceleration effect depends on the ratio between \(B_{r0}\) and \(\partial_{t}B_{z0}\), but it does not depend on the magnitude of \(B_{z0}\) itself. So, there is no need to create a large current  
Figure 4: The characteristic accumulation time \(\tau_{n}\) as a function of \(\beta=v/c\) for different values of \(\Gamma\). Curve 1 corresponds \(\Gamma=10^{-4}\); curve 2 corresponds \(\Gamma=10^{-6}\); and curve 3 corresponds \(\Gamma=10^{-8}\). Figure 5: Schematic of acceleration section: (**a**) accelerator design and (**b**) acceleration zone with plasma containing electrons, dust particles, and ions. in the solenoid, and one can expect to achieve significant acceleration of the plasma flow. Unfortunately, the described method of creating the required configuration of magnetic fields does not seem optimal since the ideal field distribution presented in Figure 5b can hardly be fully realized in the scheme shown in Figure 5a. However, one can say that this is a solvable issue from the technical and physical point of view, and we hope to discuss this question later. In this regard, we consider the energy losses associated with these factors via the parameter \(\zeta\), changes in which have a relatively weak effect on the value \(t_{m}\). For example, by varying the \(D\) value in Equation (20), we can estimate the characteristic time \(t_{m}\) with respect to the conceivable characteristics of a spacecraft actually used for near-Earth flight. We shall proceed using information about Apollo 11, the first spacecraft that made such a flight. As is known [2], the mass of Apollo 11 was 15 tons, of which 10 tons were fuel. In our estimations, we take the characteristic spacecraft mass from the range \(10^{4}\leq m\leq 10^{5}\) kg.