This paper focuses on the operational details of the Variable Specific Impulse Magnetoplasma Rocket (abbreviated as VASIMR), which is basically a rocket using plasma as its mode of propulsion for accelerating itself into space over the conventional chemical propulsion rockets. The VASIMR is an electro-magnetic thruster. It focuses on bridging the gap between high thrust-low specific impulse and low speed-high specific impulse characteristics, the two important parameters which define the performance of a rocket. These parameters can be altered for a VASIMR as per the mission requirements by altering the amount of power supplied. Such rockets are manufactured by the Ad Astra Rocket Company whose headquarters are based in Houston, Texas, United States.
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The engine consists of a covering made up of electromagnets, and a magnetic field links the three stages. The VASIMR is advantageous over other rockets because its life expectancy is increased due to the absence of physical electrodes and because every part of the engine is magnetically shielded from the plasma. Thus, the erosion of electrodes is avoided as is the wearing of the engine components. The thrust to specific impulse ratio can be controlled by selectively portioning the Radio Frequency (RF) power to the helicon and ICH couplers, along with the proper adjustment of propellant flow.
In the missions where fuel consumption or the specific impulse is a major factor for consideration, the Solar Electric Propulsion (SEP) method is employed. As VASIMR develops high power, solar energy can be effectively used for reducing the fuel requirements for transportation in space.SEP can be utilized for the following missions
The VASIMR method for heating plasma was originally developed during nuclear fusion research. VASIMR is intended to bridge the gap between high thrust, low specific impulse chemical rockets and low thrust, high specific impulse electric propulsion, but has not yet demonstrated high thrust. The VASIMR concept originated in 1977 with former NASA astronaut Franklin Chang Díaz, who has been developing the technology ever since.[2]
Over the past few decades various types of electric propulsion devices have been developed and successfully utilized in space missions, e.g., ion-gridded thrusters in DEEP-SPACE 1 (Brophy 2002) and HAYABUSA/MUSES-C missions (Kuninaka et al. 2006), a Hall thruster in SMART 1 (Koppel et al. 2005) mission, and so on. Representative important parameters showing the propulsion performance are a thrust F, a specific impulse \(I_\text sp\), and a thruster efficiency \(\eta\), where the latter two can be given as
with the mass flow rate \(\dotm\) of the propellant gas, the gravitational acceleration g, and the electric power P. Since the ionized propellant is accelerated via hydrodynamic, electrostatic, and electromagnetic acceleration processes induced by an electric power obtained in space, the electric power can be converted into the material momentum in the electric propulsion devices, yielding higher specific impulse and reducing the propellant mass mounted on the system. Hence the dynamics of the ionized gas, i.e., the plasmas, significantly affect the thruster performance. As seen in Eqs. (1) and (2), the specific impulse \(I_\text sp\) and the thruster efficiency \(\eta\) can be assessed by measuring the thrust force F with the given mass flow rate \(\dotm\) of the propellant and the electric power P. Therefore, the direct measurement of the thrust is the most important experimental issue for the thruster assessment.
Schematic diagram of a the variable specific impulse magnetoplasma rocket (VASIMR), b the inductively-coupled plasma (ICP) thruster, c the helicon plasma thruster, and d the electron cyclotron resonance (ECR) plasma thruster
Recently measured thrust and the simply calculated specific impulse and thruster efficiency from the mass flow rate of the propellant and the rf generator power. The thruster assessed here has the 95-mm-inner-diameter source tube and the gas injection near the thruster exit (Takahashi and Ando 2018)
Over the past several years, several research groups have performed the thrust assessment of the rf magnetic nozzle plasma thrusters using the thrust balances or using the target techniques as briefly summarized in Table 1. The specific impulse \(I_\text sp\), the thrust-to-power ratio \(F/P_\text rf\), and the thruster efficiency \(\eta _\text rf\) are also calculated from the measured thrust F and the mass flow rate of the propellant, where the rf generator output power \(P_\text rf\) is used for the calculation. Although the first two experiments in 2011 showed the very poor thruster efficiency less than a percent, it can be recognized that the thruster performance is gradually improved. Figure 22 shows the most recently measured thrust F, the specific impulse \(I_\text sp\), and the thruster efficiency \(\eta _\text rf\) calculated using the rf power \(P_\text rf\) from the generator, where the 95-mm-diameter source tube and the gas injection near the open source exit are employed (Takahashi and Ando 2018). It should be mentioned that the data in Fig. 22a present both the thrust and the specific impulse, since the specific impulse is proportional to the thrust for the constant mass flow rate of the propellant. The result clearly shows the best performance to this date and the thruster efficiency is now approaching \(\sim 20\)%. Eqs. (1) and (2) give the relation of the thruster efficiency \(\eta _\text rf\), the specific impulse \(I_\text sp\), and the thrust-to-power ratio \(F/P_\text rf\) as 2ff7e9595c
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