Propulsion

Introduction

Space propulsion is used to three things: lift the launch vehicle from the surface into low-Earth orbit (LEO), transfer payloads from LEO to higher orbits or interplanetary trajectories and provide attitude control and orbit corrections.

In this section, we are going to cover how to choose a thruster for a satellite given some performance indicators like the Delta-V. This guide is oriented towards the use of commercial off-the-shelf components instead of being a preliminary design guide of chemical or electrical thrusters (which is a very complex science), and therefore will not cover launch propulsion systems.

The most important data that the engineer in charge of the propulsion subsystem should provide to the design team at the end of each iteration are:

  • Propellant budget, detailing how much mass is reserved for orbit transfers, orbit corrections and attitude control
  • Mass of the thruster(s)
  • Power consumed by the thruster in stand-by and in service
  • Total size of the thruster(s)

Other components that are part of the propulsion subsystem are the tanks and the lines and pressure-regulating equipment that connects the tank to the thrusters. In this guide, we are going to suppose that the thruster is a component off-the-shelf and, therefore, we are not going to design that connecting system.

The propulsion subsystem has two purposes: translational control manoeuvres and attitude control manoeuvres. Depending on its purposes, its requirements and location on the spacecraft are different. For example, a translational control thruster should be aligned with the centre of mass, whereas an attitude control one should be as far as possible from it to increase the lever arm and therefore reduce the required thrust.

In terms of the impact of this subsystem to the rest of the vehicle, its power consumption is low unless it requires heated propellant or it is an electric thruster. Regarding the mass budget, its impact is big since the propellant must be burn and expelled to create thrust. Finally, thermally speaking, it depends on the type of thruster but in principle we can say that at least we should prevent the propellant and the lines from freezing, so heaters may be required.

In Table 12 are summarised the principal propulsion technologies and where they can be applied.

Table 12 Principal options for spacecraft propulsion systems, from [WL99]. P: perigee, A: apogee
Propulsion technology Orbit insertion Orbit Maintenance and Manoeuvring Attitude Control Typical Isp (s)
Cold gas PA X X 30-70
Solid PA     280-300
Liquid: monopropellant PA X X 220-240
Liquid: bipropellant PA X X 305-310
Liquid: dual mode PA X X 313-322
Hybrid PA X   250-340
Electric A X   300-3000

Design Process

  1. List applicable spacecraft propulsion functions, like orbit insertion, orbit maintenance, attitude control and controlled de-orbit or reentry.
  2. Determine ΔV budget and thrust level constraints for orbit insertion and maintenance.
  3. Determine total impulse for attitude control, thrust levels for control authority, duty cycles (% on/off, total number of cycles) and mission life requirements.
  4. Determine propulsion system options: combined or separate propulsion systems for orbit and attitude control, high vs. low thrust, liquid vs. solid vs. electric propulsion technology.
  5. Estimate key parameters for each option: effective specific impulse for orbit and attitude control, propellant mass, propellant and pressurant volume, configure the subsystem and create equipment list.
  6. Estime the total mass and power for each option.
  7. Establish baseline propulsion system.
  8. Document results and iterate as required.

The Rocket Equation

Also known as Tsiolkovsky equation (17), it is the most important equation in mission analysis and therefore it is important too for the propulsion subsystem: given a required ΔV (Delta-V, m/s) we can estimate the required propellant mass if we know the specific impulse \(I_{sp}\) (s), which is the most characteristic parameter of a thruster.

(17)\[\Delta V = I_{sp} g_0 \log{\frac{M_0}{M_f}}\]

In Equation (17) it is assumed that the manoeuvre is applied instantaneously, so it is only valid for short bursts. \(M_0\) and \(M_f\) are the pre- and post-manoeuvre masses (kg) of the vehicle, and \(g_0\) is 9.81 m/s². From this equation, we can compute the required propellant mass for our ΔV, since that would be the change in the total mass of the vehicle:

\[\Delta m_p = M_0 \left( 1 - e^{-\Delta V / I_{sp} g_0}\right)\]

Types of Rockets

This is a brief description of the types of rockets or propulsion systems that we saw in Table 12. The principal propulsion technologies are: cold gas, chemical and electrical.

Cold gas

Cold gas propulsion is just a controlled, pressurised gas source and a nozzle. There is no combustion and there is only a gas expelled. Its the simplest form of a rocket engine.

Chemical

Chemical thrusters work redirecting the resultant gases from the combustion of the propellants. Depending on the nature of the propellants, we can have liquid or solid chemical propulsion, with important changes in the internal architecture of the thrusters. The internal architecture of the thrusters is a topic that is not going to be covered due to its complexity and lack of interest when using COTS.

The terminology solid, liquid or hybrid refers to the initial state of the stored propellant.

Liquid rocket systems use liquid propellants that are fed to the combustion chamber by gas pressurisation or a pump. Depending on the number of components, these rockets are called bipropellant if they use two propellants or monopropellant if only one.

On the one hand, bipropellant systems are more complex since they usually have a fuel and an oxidiser that chemically react in a combustion process, but they can provide a higher specific impulse. On the other hand, monopropellant systems are simpler and therefore more reliable; the propellant that they use can ignite and provide the required energy without more components. A very common monopropellant is hydrazine, since it is stable under normal storage conditions, has a clean decomposition process and is very easy to handle, although it must be done with care due to its toxicity. It is the most common type of propulsion for spacecraft attitude and velocity control.

Hybrid systems have the propellants in different states, usually the fuel is a solid and the oxidiser is a gas or a liquid. They are not very common.

Finally, solid rockets have their propellant in solid form. They have lower performance than liquid rockets but they are simpler. Their main limitation is that, once they have been ignited, they cannot be shut down, that is to say, they can only be used once.

Electric

Electric propulsion accelerates the working fluid using electrical power obtained from the Sun as solar energy, nuclear or thermal. This acceleration is the responsible of the creation of thrust once the fluid is expelled. There are different ways to use electric power to create this thrust, and depending on the way we can establish three main classes of electric propulsion systems: electrothermal, electrostatic and electromagnetic.

Since these propulsion systems rely heavily on electric power and propellant mass, it is important to measure their efficiency to make a good comparison. The relationship between the power \(P\) (W), the thrust \(F\) (N), the specific impulse \(I_{sp}\), the mass flow rate \(\dot{m}\) (kg/s) and the efficiency \(\eta\) is:

\[P = \frac{F^2}{2 \dot{m}\eta} = \frac{F I_{sp} g}{2 \eta}\]
Table 13 Main characteristics of electric propulsion systems
Class Name Propellant Power Specific impulse (s) Efficiency Thrust
Electrothermal Resistojet Hydrazine, ammoniac 0.5-1.5 kW 300 80% 0.1-0.5N
Electrothermal Arcjet Hydrazine, hydrogen 0.3-100 kW 500-2,000 35% 0.2-2N
Electrostatic Ion Xenon 0.5-2.5 kW 3,000 60%-80% 10-200mN
Electrostatic Hall Xenon 1.5-5kW 1,500-2,000 50% 80-200mN
Electrostatic FEED Indium, caesium 10-150W 6,000-10,000 30%-90% 0.001-2mN
Electrostatic Coloidal Glycerine 5-50 W 500-1,500 (-) 0.001-1mN
Electromagnetic PPT Teflón 1-200 W 1,000 5% 1-100mN
Electromagnetic MPD Hydrogen, ammoniac, lithium 1-4000kW 2,000-5,000 25% 1-200N
Electromagnetic VASIMR Hydrogen 1-10 MW 3,000-30,000 20%-60% 1-2kN

References and Other Resources