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# PowerPedia:Spacecraft propulsion

Spacecraft propulsion is used to change the velocity of spacecraft and artificial satellites, or in short, to provide delta-v. There are many different methods. Each method has drawbacks and advantages, and spacecraft propulsion is an active area of research. Most spacecraft today are propelled by heating the reaction mass and allowing it to flow out the back of the vehicle. This sort of engine is called a rocket engine.

All current spacecraft use chemical rocket (bipropellant or solid-fuel) for launch, though some (such as the Pegasus rocket and SpaceShipOne) have used air-breathing engines on their first stage. Most satellites have simple reliable chemical rockets (often monopropellant rockets) or resistojet rockets to keep their station, although some use momentum wheels for attitude control. Newer geo-orbiting spacecraft are starting to use electric propulsion for north-south stationkeeping. Interplanetary vehicles mostly use chemical rockets as well, although a few have experimentally used ion thrusters with some success (a form of electric propulsion).

## The necessity for propulsion systems

Artificial satellites must be launched into orbit, and once there they must be placed in their nominal orbit. Once in the desired orbit, they often need some form of attitude control so that they are correctly pointed with respect to the Earth, the Sun, and possibly some astronomical object of interest. They are also subject to drag from the thin atmosphere, so that to stay in orbit for a long period of time some form of propulsion is occasionally necessary to make small corrections (orbital stationkeeping). Many satellites need to be moved from one orbit to another from time to time, and this also requires propulsion. When a satellite has exhausted its ability to adjust its orbit, its useful life is over.

Spacecraft designed to travel further also need propulsion methods. They need to be launched out of the Earth's atmosphere just as satellites do. Once there, they need to leave orbit and move around.

For interplanetary travel, a spacecraft must use its engines to leave Earth orbit. Once it has done so, it must somehow make its way to its destination. Current interplanetary spacecraft do this with a series of short-term orbital adjustments. In between these adjustments, the spacecraft simply falls freely along its orbit. The simplest fuel-efficient means to move from one circular orbit to another is with a Hohmann transfer orbit: the spacecraft begins in a roughly circular orbit around the Sun. A short period of thrust in the direction of motion accelerates or decelerates the spacecraft into an elliptical orbit around the Sun which is tangential to its previous orbit and also to the orbit of its destination. The spacecraft falls freely along this elliptical orbit until it reaches its destination, where another short period of thrust accelerates or decelerates it to match the orbit of its destination. Special methods such as aerobraking are sometimes used for this final orbital adjustment.

Some spacecraft propulsion methods such as solar sails provide very low but inexhaustible thrust; an interplanetary vehicle using one of these methods would follow a rather different trajectory, either constantly thrusting against its direction of motion in order to decrease its distance from the Sun or constantly thrusting along its direction of motion to increase its distance from the Sun.

Spacecraft for interstellar travel also need propulsion methods. No such spacecraft has yet been built, but many designs have been discussed. Since interstellar distances are very great, a tremendous velocity is needed to get a spacecraft to its destination in a reasonable amount of time. Acquiring such a velocity on launch and getting rid of it on arrival will be a formidable challenge for spacecraft designers.

## Effectiveness of propulsion systems

When in space, the purpose of a propulsion system is to change the velocity v of a spacecraft. Since this is more difficult for more massive spacecraft, designers generally discuss momentum, mv. The amount of change in momentum is called impulse. So the goal of a propulsion method in space is to create an impulse.

When launching a spacecraft from the Earth, a propulsion method must overcome a higher gravitational pull to provide a net positive acceleration. In orbit, the spacecraft tangential velocity provides a centrifugal force that counterweighs the gravity pull at a given path (which is actually the orbit path) so that any additional impulse even very tiny will result in a change in the orbit path.

The rate of change of velocity is called acceleration, and the rate of change of momentum is called force. To reach a given velocity, one can apply a small acceleration over a long period of time, or one can apply a large acceleration over a short time. Similarly, one can achieve a given impulse with a large force over a short time or a small force over a long time. This means that for maneuvering in space, a propulsion method that produces tiny accelerations but runs for a long time can produce the same impulse as a propulsion method that produces large accelerations for a short time. When launching from a planet, tiny accelerations cannot overcome the planet's gravitational pull and so cannot be used.

The law of conservation of momentum means that in order for a propulsion method to change the momentum of a space craft it must change the momentum of something else as well. A few designs take advantage of things like magnetic fields or light pressure in order to change the spacecraft's momentum, but in free space the rocket must bring along some mass to accelerate away in order to push itself forward. Such mass is called reaction mass.

In order for a rocket to work, it needs two things: reaction mass and energy. The impulse provided by launching a particle of reaction mass having mass m at velocity v is mv. But this particle has kinetic energy mv2/2, which must come from somewhere. In a conventional solid fuel rocket, the fuel is burned, providing the energy, and the reaction products are allowed to flow out the back, providing the reaction mass. In an ion thruster, electricity is used to accelerate ions out the back. Here some other source must provide the electrical energy (perhaps a solar panel or a nuclear reactor) while the ions provide the reaction mass.

When discussing the efficiency of a propulsion system, designers often focus on effectively using the reaction mass. Reaction mass must be carried along with the rocket and is irretrievably consumed when used. One way of measuring the amount of impulse that can be obtained from a fixed amount of reaction mass is the specific impulse, the impulse per unit weight-on-Earth (typically designated by Isp). The unit for this value is seconds. Since the weight on Earth of the reaction mass is often unimportant when discussing vehicles in space, specific impulse can also be discussed in terms of impulse per unit mass. This alternate form of specific impulse uses the same units as velocity (e.g. m/s), and in fact it is equal to the effective exhaust velocity of the engine (typically designated ve). Confusingly, both values are sometimes called specific impulse. The two values differ by a factor of g, the acceleration due to gravity on the Earth's surface (Ispg = ve).

A rocket with a high exhaust velocity can achieve the same impulse with less reaction mass. However, the energy required for that impulse is proportional to the square of the exhaust velocity, so that more mass-efficient engines require much more energy. This is a problem if the engine is to provide a large amount of thrust. To generate a large amount of impulse per second, it must use a large amount of energy per second. So highly efficient engines require enormous amounts of energy per second to produce high thrusts. As a result, most high-efficiency engine designs also provide very low thrust.

### Calculations

Burning the entire usable propellant of a spacecraft through the engines in a straight line in free space would produce a net velocity change to the vehicle; this number is termed 'delta-v'.

The total Δv of a vehicle can be calculated using the rocket equation, where M is the mass of fuel (or rather the mass of propellant), P is the mass of the payload (including the rocket structure), and ve is the velocity of the rocket exhaust. This is known as the Tsiolkovsky rocket equation:

$\Delta V = -v_e \ln \left(\frac{P}{M+P}\right)$

For historical reasons, as discussed above, ve is sometimes written as

ve = Ispgo

where Isp is the specific impulse of the rocket, measured in seconds, and go is the gravitational acceleration at sea level.

For a long voyage, the majority of the spacecraft's mass may be reaction mass. Since a rocket must carry all its reaction mass with it, most of the first reaction mass goes towards accelerating reaction mass rather than payload. If we have a payload of mass P, the spacecraft needs to change its velocity by Δv, and the rocket engine has exhaust velocity ve, then the mass M of reaction mass which is needed can be calculated using the rocket equation and the formula for Isp

$M = P \left(e^{\Delta v/v_e}-1\right)$

For Δv much smaller than ve, this equation is roughly linear, and little reaction mass is needed. If Δv is comparable to ve, then there needs to be about twice as much fuel as combined payload and structure (which includes engines, fuel tanks, and so on). Beyond this, the growth is exponential; speeds much higher than the exhaust velocity require very high ratios of fuel mass to payload and structural mass.

In order to achieve this, some amount of energy must go into accelerating the reaction mass. Every engine will waste some energy, but even assuming 100% efficiency, the engine will need energy amounting to

$\frac{1}{2} Mv_e^2$

Comparing the rocket equation (which shows how much energy ends up in the final vehicle) and the above equation (which shows the total energy required) shows that even with 100% engine efficiency, certainly not all energy supplied ends up in the vehicle - some of it, indeed usually most of it, ends up as kinetic energy of the exhaust.

For a mission, for example, when launching from or landing on a planet, the effects of gravitational attraction and any atmospheric drag must be overcome by using fuel. It is typical to combine the effects of these and other effects into an effective mission delta-v. For example a launch mission to low Earth orbit requires about 9.3-10 km/s delta-v. These mission delta-vs are typically numerically integrated on a computer.

Suppose we want to send a 10,000 kg space probe to Mars. The required Δv from LEO is approximately 3000 m/s, using a Hohmann transfer orbit. (A manned probe would need to take a faster route and use more fuel). For the sake of argument, let us say that the following thrusters may be used:

 Engine Effective Exhaust Velocity(m/s) Specific impulse(s) Fuel mass(kg) Energy required(GJ) Energy per kgof propellant minimum powerper N of thrust Solid rocket 1,000 100 190,000 95 500 kJ 0.5 kW Bipropellant rocket 5,000 500 8,200 103 12.6 MJ 2.5 kW Ion thruster 50,000 5,000 620 775 1.25 GJ 25 kW VASIMR 300,000 30,000 100 4,500 45 GJ 150 kW

Observe that the more fuel-efficient engines can use far less fuel; its mass is almost negligible (relative to the mass of the payload and the engine itself) for some of the engines. However, note also that these require a large total amount of energy. For earth launch engines require a thrust to weight ratio of much more than unity. To do this they would have to be supplied with Gigawatts of power — equivalent to a major metropolitan generating station. This would need to be carried on the vehicle, which is clearly impractical.

Instead, a much smaller, less powerful generator may be included which will take much longer to generate the total energy needed. This lower power is only sufficient to accelerate a tiny amount of fuel per second, but over long periods the velocity will be finally achieved. For example. it took the Smart 1 more than a year to reach the Moon, while with a chemical rocket it takes a few days. Because the ion drive needs much less fuel, the total launched mass is usually lower, which typically results in a lower overall cost.

Interestingly, for a mission delta-v, there is a fixed Isp that minimises the overall energy used by the rocket. This comes to an exhaust velocity of about ⅔ of the mission delta-v (see also the energy computed from the rocket equation). Drives such as VASIMR, and to a lesser extent other Ion thrusters have exhaust velocities that can be enormously higher than this ideal, and thus end up powersource limited and give very low thrust. Where the vehicle performance is power limited, e.g. if solar power or nuclear power is used, then in the case of a large ve the maximum acceleration is inversely proportional to it. Hence the time to reach a required delta-v is proportional to ve. Thus the latter should not be too large.

## Propulsion methods

Propulsion methods can be classified based on their means of accelerating the reaction mass. There are also some special methods for launches, planetary arrivals, and landings.

### Rocket engines

A rocket engine is a reaction engine that can be used for spacecraft propulsion as well as terrestrial uses, such as missiles. Rocket engines take their reaction mass from one or more tanks and form it into a jet, obtaining thrust in accordance with Newton's third law. Most rocket engines are internal combustion engines, although non combusting forms exist.

Most rocket engines are internal combustion heat engines (although non combusting forms exist). Rocket engines generally produce a high temperature reaction mass, as a hot gas. This is achieved by combusting a liquid or gaseous fuel with an oxidiser within a combustion chamber. The extremely hot gas is then allowed to escape through a high-expansion ratio nozzle. This bell-shaped nozzle is what gives a rocket engine its characteristic shape. The effect of the nozzle is to dramatically accelerate the mass, converting most of the thermal energy into kinetic energy. Exhaust speeds as high as 10 times the speed of sound at sea level are not uncommon.

Rockets emitting plasma can potentially carry out reactions inside a magnetic bottle and release the plasma via a magnetic nozzle, so that no solid matter need come in contact with the plasma. Of course, the machinery to do this is complex, but research into nuclear fusion has developed methods, some of which have been used in speculative propulsion systems.

Classic rocket engines produce a high temperature, hypersonic gaseous exhaust. This is achieved by the combustion of solid, liquid or gaseous propellant, containing oxidiser and a fuel, within a combustion chamber at high pressure. The hot gas produced is then allowed to escape through a narrow hole (the 'throat'), into a high-expansion ratio nozzle. The effect of the nozzle is to dramatically accelerate the mass, converting most of the thermal energy into kinetic energy. The large bell or cone shaped expansion nozzle gives a rocket engine its characteristic shape. Exhaust speeds as high as 10 times the speed of sound at sea level are not uncommon.

Part of the rocket engine's thrust comes from the gas pressure inside the combustion chamber but the majority comes from the pressure against the inside of the expansion nozzle. Inside the combustion chamber the gas produces a similar force against all the sides of the combustion chamber but the throat gives no force producing an unopposed resultant force from the diametrically opposite end of the chamber. As the gases (adiabatically) expand inside the nozzle they press against the bell's walls forcing the rocket engine in one direction, and accelerating the gases in the opposite direction.

For optimum performance hot gas is used because it maximises the speed of sound at the throat-for aerodynamic reasons the flow goes sonic ("chokes") at the throat, so the highest speed there is desirable. By comparison, at room temperature the speed of sound in air is about 340m/s, the speed of sound in the hot gas of a rocket engine can be over 1700m/s.

The expansion part of the rocket nozzle then multiplies the speed of the flow by a further factor, typically between 1.5 and 4 times, giving a highly collimated exhaust jet. The speed ratio of a rocket nozzle is mostly determined by its area expansion ratio—the ratio of the area of the throat to the area at the exit, but details of the gas properties are also important. Larger ratio nozzles are more massive and bulkier, but they are able to extract more heat from the combustion gases, which become lower in pressure and colder, but also faster.

A significant complication arises when launching a vehicle from the Earth's surface as the ambient atmospheric pressure changes with altitude. For maximum performance it turns out that the pressure of the gas leaving a rocket nozzle should be the same as ambient pressure; if lower the vehicle will be slowed by the difference in pressure between the top of the engine and the exit, if higher then this represents pressure that the bell has not turned into thrust. To achieve this ideal, the diameter of the nozzle would need to increase with altitude, which is difficult to arrange. A compromise nozzle is generally used and some percentage reduction in performance occurs. To improve on this, various exotic nozzle designs such as the plug nozzle, stepped nozzles, the expanding nozzle and the aerospike have been proposed, each having some way to adapt to changing ambient air pressure and each allowing the gas to expand further against the nozzle giving extra thrust at higher altitude.

### Airbreathing engines for launch

Studies generally show that conventional air-breathing engines, such as ramjets or turbojets are basically too heavy (have too low a thrust/weight ratio) to give any significant performance improvement when installed on a launch vehicle. However, they can be used on a separate lift vehicle (e.g. X-1, Pegasus and SS1). On the other hand, very lightweight or very high speed engines have been proposed that take advantage of the air during ascent:

A jet engine is an engine that discharges a fast moving jet of fluid to generate thrust in accordance with Newton's third law of motion. This broad definition of jet engines includes turbojets, turbofans, rockets and ramjets and water jets, but in common usage, the term generally refers to a gas turbine used to produce a jet of high speed exhaust gases for special propulsive purposes.

Engines that may need to operate at low hypersonic speeds could theoretically have much higher performance if a heat exchanger is used to cool the incoming air. The low temperature allows lighter materials to be used and combustors run at their maximum speeds (ordinarily, fuel flow must be reduced to prevent the turbines from melting, but doing so greatly reduces thrust) This leads to plausible designs like SABRE, that might permit single-stage-to-orbit. The Skylon Spaceplane, and ATREX, that might permit jet engines to be used as boosters for space vehicles.

### Electromagnetic acceleration of reaction mass

Rather than relying on high temperature and fluid dynamics to accelerate the reaction mass to high speeds, there are a variety of methods that use electrostatic or electromagnetic forces to accelerate the reaction mass directly. Usually the reaction mass is a stream of ions. Such an engine requires electric power to run, and high exhaust velocities require large amounts of energy.

For these drives it turns out that to a reasonable approximation fuel use, energetic efficiency and thrust are all inversely proportional to exhaust velocity. Their very high exhaust velocity means they require huge amounts of energy and thus with practical powersources provide low thrust, but use hardly any fuel.

For some missions, solar energy may be sufficient, and has very often been used, but for others nuclear energy will be necessary; engines drawing their power from a nuclear source are called nuclear electric rockets.

With any current source of power, chemical, nuclear or solar, the maximum amount of power that can be generated greatly limits the maximum amount of thrust that can be produced to a small value. Power generation also adds significant mass to the spacecraft, and ultimately the weight of the power source limits the performance of the vehicle. Current nuclear power generators are approximately half the weight of solar panels per watt of energy supplied, at terrestrial distances from the Sun. Chemical power generators are not used due to the far lower total available energy. Beamed power to the spacecraft shows potential.

The dissipation of waste heat from the powerplant may make any propulsion system requiring a separate power source infeasible for interstellar travel.

Some electromagnetic methods:

### Systems without reaction mass carried within the spacecraft

The law of conservation of momentum states that any engine which uses no reaction mass cannot move the center of mass of a spaceship (changing orientation, on the other hand, is possible). But space is not empty, especially space inside the Solar System; there are gravitation fields, magnetic fields, solar wind and solar radiation. Various propulsion methods try to take advantage of these. However, since these phenomena are diffuse in nature, corresponding propulsion structures need to be proportionately large.

Space drives that need no (or little) reaction mass:

For changing the orientation of a satellite or other space vehicle, conservation of angular momentum does not pose a similar constraint. Thus many satellites use momentum wheels to control their orientations. These cannot be the only system for controlling satellite orientation, as the angular momentum built up due to torques from external forces such as solar, magnetic or tidal forces eventually needs to be "bled off" using a secondary system.

### Launch mechanisms

High thrust is of vital importance for Earth launch, thrust has to be greater than weight (see also gravity drag). Many of the propulsion methods above give a thrust/weight ratio of much less than 1, and so cannot be used for launch.

Exhaust toxicity or other side effects can also have detrimental effects on the environment the spacecraft is launching from, ruling out other propulsion methods, such as most nuclear engines, at least for use from the Earths surface.

Therefore, all current spacecraft use chemical rocket engines (bipropellant or solid-fuel) for launch.

One advantage that spacecraft have in launch is the availability of infrastructure on the ground to assist them. Proposed ground-assisted launch mechanisms include:

### Planetary arrival and landing

When a vehicle is to enter orbit around its destination planet, or when it is to land, it must adjust its velocity. This can be done using all the methods listed above (provided they can generate a high enough thrust), but there are a few methods that can take advantage of planetary atmospheres and/or surfaces.

• Aerobraking allows a spacecraft to reduce the high point of an elliptical orbit by repeated brushes with the atmosphere at the low point of the orbit. This can save a considerable amount of fuel since it takes much less delta-V to enter an elliptical orbit compared to a low circular orbit. Since the braking is done over the course of many orbits, heating is comparatively minor, and a heat shield is not required. This has been done on several Mars missions such as Mars Global Surveyor, Mars Odyssey and Mars Reconnaissance Orbiter, and at least one Venus mission, Magellan.
• Aerocapture is a much more aggressive manoeuver, converting an incoming hyperbolic orbit to an elliptical orbit in one pass. This requires a heat shield and much trickier navigation, since it must be completed in one pass through the atmosphere, and unlike aerobraking no preview of the atmosphere is possible. If the intent is to remain in orbit, then at least one more propulsive maneuver is required after aerocapture—otherwise the low point of the resulting orbit will remain in the atmosphere, resulting in eventual re-entry. Aerocapture has not yet been tried on a planetary mission, but the re-entry skip by Zond 6 and Zond 7 upon lunar return were aerocapture maneuvers, since they turned a hyperbolic orbit into an elliptical orbit. On these missions, since there was no attempt to raise the perigee after the aerocapture, the resulting orbit still intersected the atmosphere, and re-entry occurred at the next perigee.
• Parachutes can land a probe on a planet with an atmosphere, usually after the atmosphere has scrubbed off most of the velocity, using a heat shield.
• Airbags can soften the final landing.
• Lithobraking, or stopping by simply smashing into the target, is usually done by accident. However, it may be done deliberately with the probe expected to survive (see, for example, Deep Space 2). Very sturdy probes and low approach velocities are required.

Gravitational slingshots can also be used to carry a probe onward to other destinations.

### Methods requiring new principles of physics

In addition, a variety of hypothetical propulsion techniques have been considered that would require entirely new principles of physics to realize. To date, such methods are highly speculative and include:

## Table of methods and their specific impulse

Below is a summary of some of the more popular, proven technologies, followed by increasingly speculative methods.

Three numbers are shown. The first is the effective exhaust velocity: the equivalent speed that the propellant leaves the vehicle. This is not necessarily the most important characteristic of the propulsion method, thrust and power consumption and other factors can be, however:

• if the delta-v is much more than the exhaust velocity, then exorbitant amounts of fuel are necessary (see the section on calculations, above)
• if it is much more than the delta-v, then, proportionally more energy is needed; if the power is limited, as with solar energy, this means that the journey takes a proportionally longer time

The second and third are the typical amounts of thrust and the typical burn times of the method. Outside a gravitational potential small amounts of thrust applied over a long period will give the same effect as large amounts of thrust over a short period. (This result does not apply when the object is significantly influenced by gravity.)

Propulsion methods
Method Effective Exhaust Velocity
(m/s)
Thrust
(N)
Duration
Propulsion methods in current use
Solid rocket 1,000 - 4,000 103 - 107 minutes
Hybrid rocket 1,500 - 4,200 minutes
Monopropellant rocket 1,000 - 3,000 0.1 - 100 milliseconds - minutes
Bipropellant rocket 1,000 - 4,700 0.1 - 107 minutes
Tripropellant rocket 2,500 - 4,500 minutes
Resistojet rocket 2,000 - 6,000 10-2 - 10 minutes
Arcjet rocket 4,000 - 12,000 10-2 - 10 minutes
Hall effect thruster (HET) 8,000 - 50,000 10-3 - 10 months
Electrostatic ion thruster 15,000 - 80,000 10-3 - 10 months
Field Emission Electric Propulsion (FEEP) 100,000 - 130,000 10-6 - 10-3 weeks
Magnetoplasmadynamic thruster (MPD) 20,000 - 100,000 100 weeks
Pulsed plasma thruster (PPT)
Pulsed inductive thruster (PIT) 50,000 20 months
Nuclear electric rocket As electric propulsion method used
Tether propulsion N/A 1 - 1012 minutes
Currently feasible propulsion methods
Solar sails N/A 9 per km²
(at 1 AU)
Indefinite
Mass drivers (for propulsion) 30,000 - ? 104 - 108 months
Orion Project (Near term nuclear pulse propulsion) 20,000 - 100,000 109 - 1012 several days
Variable specific impulse magnetoplasma rocket (VASIMR) 10,000 - 300,000 40 - 1,200 days - months
Nuclear thermal rocket 9,000 105 minutes
Solar thermal rocket 7,000 - 12,000 1 - 100 weeks
Air-augmented rocket 5,000 - 6,000 seconds-minutes
Liquid air cycle engine 4,500 seconds-minutes
SABRE 30,000/4,500 minutes
Dual mode propulsion rocket
Technologies requiring further research
Magnetic sails N/A Indefinite Indefinite
Mini-magnetospheric plasma propulsion 200,000 ~1 N/kW months
Nuclear pulse propulsion (Project Daedalus' drive) 20,000 - 1,000,000 109 - 1012 half hour
Gas core reactor rocket 10,000 - 20,000 103 - 106
Antimatter catalyzed nuclear pulse propulsion 20,000 - 400,000 days-weeks
Nuclear salt-water rocket 100,000 103 - 107 half hour
Fission sail
Fission-fragment rocket 10,000,000
Nuclear photonic rocket 300,000,000 10-5 - 1 years-decades
Significantly beyond current engineering
Fusion rocket 1,300,000-36,000,000
Bussard ramjet
Antimatter rocket 10,000,000-100,000,000
Redshift rocket
gravitoelectromagnetic toroidal launchers

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