Introduction to Astronautics Sissejuhatus kosmonautikasse

Tallinn University of Technology Introduction to AstronauticsSissejuhatus kosmonautikasse Vladislav Pustõnski 2009

Nuclear propulsion Non-chemical propulsion methods • Nuclear rocket propulsion includes a variety of methods where (thermo)nuclear reactions provides an energy source. The following methods may be pointed out: • thermal propulsion: nuclear reactor heats a working fluid (usually hydrogen, but other fluids are also possible) which is ejected and creates thrust; • pulse propulsion: small nuclear explosions behind a protective plate accelerate the rocket from outside; • radioisotope propulsion: decay of radioactive isotopes provide energy to heat a working fluid; • nuclear electric propulsion: nuclear reactions are used to produce electricity supply for one of the electric propulsion methods; • variations of these methods and some other original applications. • The greatest advantage of the nuclear propulsion that it enables to get quite high values of the specific impulse that are not achievable with chemical propulsion methods. That means that the rocket device may be much smaller, with much less working fluid onboard. Some nuclear propulsion methods require quite simple engines that may be more reliable than chemical engines. • The greatest disadvantage of nuclear propulsion is its fuel. Nuclear fuel is not safe due to high radioactivity, and its production and handling are very risky. A launch failure of a rocket with a nuclear powerplant onboard (or carrying a payload with a nuclear device) may have grave consequences. In addition, nuclear devices onboard may be potentially hazardous for

crews of manned spacecrafts. This is the reason why nuclear propulsion have not been applied so far, although most of technical problems have been solved with several propulsion methods, and the developments in this area are quite advanced. Let us look in detail the principle characteristics of the nuclear propulsion methods. Thermal propulsion The general idea of thermal propulsion is quite similar to the common thermal propulsion: the working fluid is heated and gasified in the thrust chamber, and the gases are accelerated in the nozzle giving thrust to the rocket; but nuclear reactions are used for heating, not chemical reactions. LH2 is proposed as the best choice of propellant due to its low molar weight; however, far from the Earth, where use of cryogenic propellant is problematic, other liquids may be used, among which is water. In general, the weight of nuclear reactors is large, so the thrust-to-weight ratio of nuclear thermal engines is low. The propellant tanks, the structure and the upper stages would render it smaller than unity. So it is clear that these engines cannot be used on first stages of rockets. However, they may be used on upper stages, on satellites and space probes, when the stage with the reactor need not overcome the gravity of the Earth. Use on planets with low gravity and the Moon is also possible. The most obvious and the simplest design is the solid core design, where a conventional (albeit light-weight) nuclear reactor with a solid core runs at high temperatures, and the working liquid flows through the core. In this design the specific impulse is limited by the melting temperature of the reactor elements. Nowadays this temperature is ~39000C for Hafnium carbide. With hydrogen propellant, that would provide Isp ~ 800  1000 sec. This is a very high value not achievable with any chemical propellant. Ispof LH2/LOX is as twice

as smaller. For example, a Moon land and return mission would require a mass ratio about 6-8 on LEO (characteristic velocity ~ 8 km/sec) if performed using chemical propellants; a solid core nuclear engine would make it possible to reduce the mass ratio to ~2.5. So the mass of the payload would increase ~3 times. Some designs (using of non-rigid fuel elements) would enable to increase Ispto values > 1000 sec, but the complexity would raise significantly. To overcome the temperature limitations of the solid core design, it is theoretically possible to let the fuel elements to work at higher temperatures, i.e. to melt. This is the liquid core design. Melted fuel is mixed with the propellant and is either trapped in the engine by centrifugal forces (so, the liquid rotates at high speed) either escapes with the gas (which would mean extremely radioactive exhaust and thus, such engine could be used only far from the Earth). Specific impulse values of Isp ~ 1400sec would be achievable. Gas core offers even higher Isp values, from ~ 2000 sec to ~ 5000 sec. The fuel is gasified and works at temperatures of several tens of thousands of Kelvins. It may either mix with the working fluid (and escape with it) either be separated by a transparent quartz container (which would make the cycle closed and the rocket reusable). Solid core nuclear engines were being developed in US and the USSR. In US it was the NERVA (Nuclear Engine for Rocket Vehicle Application) project to replace the upper stage of the Saturn V. Some experimental engines were built and tested. However several technical problems remained unsolved when the project was closed in the beginning of 1970th. Other studies were also performed. In the USSR it was the test engine RD-01410. Risks of launching nuclear reactors is obvious, but there are reasons to think that in the case of a launch failure a construction capable to withstand very high working temperatures and loads would not disintegrate into small parts. So it will not contaminate large areas but the remains would fall as one piece or small number of pieces contaminating a confined zone.

Pulse propulsion The general idea of the pulse propulsion is to use (thermo)nuclear explosions to accelerate the vehicle. Small bombs (in different designs their TNT equivalent may vary from tens of tons to tens of kilotons and higher) are exploded behind a steel pusher plate in the backside of the vehicle. The nuclear core of each bomb is surrounded by a specially formed layer of reaction mass (tungsten, polyethylene or some other), being a shaped charge. The directed blast of the evaporated reaction mass hits the pusher plate and accelerates the vehicle. To reduce g-forces, a complex system of shock absorbers between the pusher plate and the vehicle is installed. Each blast may add 10 – 20 m/sec to the velocity of the vehicle, so ~1000 bombs are needed to put the vehicle to LEO. The impressive specific impulse of ~ 2000  3000 sec is achievable with such design, coupled with very high thrust (which is defined only by TNT equivalent of the bombs used). So, spacecrafts with masses of many thousands of tons are possible. A lunar base may be built with only one launch, travels to distant planets could be realized within a period of one year. Most of the problems seems to be solvable with the current technology, including erosion of the pusher plate (covered with oil ablation, it would withstand numerous blasts without major erosion). The most serious problem is radioactive fallout, which is not avoidable, so use of such a vehicle would lead to massive contamination. So, the only possibility is to start it out of the atmosphere (and probably also out of the magnetosphere, to avoid trapping of exhaust radioactive materials by the magnetic field and return them to the Earth). But this means renunciation of the major advantage: cheap delivery of very large masses into space. Several projects of pulse propulsion have been studied, the most known is the project Orion, where spacecraft with masses from 300 to 8 millions of tons were examined. The project was cancelled in mid 1960th due to the problems of fallout and the cancellation of nuclear tests in space.

The basic idea of the radioisotope propulsion is to use heat of decay of radioisotopes to gasify the working fluid and to expel it from the nozzle. The difference from the thermal propulsion is use of passive decay instead of an active reactor. This design would enable specific impulse values of ~ 800 sec. The principle disadvantage of such solution is low thrust. While a nuclear reactor may produce ~ 1 GW of power, a radioisotope heat supply would produce only several kW. So, the thrust of such engine would be only several newton and it would be applicable only on small thrusters and for low-thrust propulsion. To increase the thrust, Plutonium (which has long half life and is often used for electricity supply on space probes) may be replaced by Polonium 210 which has shorter half life (~140d) and was already used in space (in heaters of the Lunokhods). Radioisotope propulsion Nuclear electric propulsion The basic idea of the nuclear electric propulsion is to use nuclear reactions to produce electricity, which powers one of the existent electric propulsion devices (see further). Conventional electric propulsion is generally of low thrust, but the thrust is limited by low power of the available electricity supply (ordinarily solar cells). If high power supply is available from a nuclear reactor, much higher thrusts would be achievable. This propulsion method may be quite promising because of high specific impulse values that electric propulsion provides.

Electric propulsion Electric propulsion methods are focused in converting electric power to thrust with the aid of working mass onboard the vehicle. The mass is accelerated by electric or electromagnetic field and ejected unidirectionally, providing the vehicle with a momentum in the opposite direction. In many applications even no nozzle is needed, since the acceleration principle is different from that of thermal rockets. Some engines work in continuous regime, other in impulse regime. Electric propulsion methods enable to achieve very high values of the specific impulse. However their thrust is very low since it is limited by power of the electricity supply onboard, and electricity is generally produced by solar cells. The low thrust of these methods confines their use to on-orbit applications (no launch vehicle may be built basing on electric propulsion methods). So, electric propulsion is used on thrusters of satellites for attitude control and station keeping. In the last years, it have been also used on space probes for low-thrust propulsion (for instance, on the Dawn asteroid probe to the Ceres and the Vesta). Electric propulsion methods are divided into three: 1) electrothermal; 2) electrostatic; 3) electromagnetic. Let us study them closer. Electrothermal methods This is the first group of methods used in practice for spacecraft acceleration. The general idea is to use electromagnetic fields for heating propellant: so, electromagnetic heating is used instead of chemical heating (as in common chemical propulsion). So, all main properties of chemical engines keep (for instance, a nozzle is needed to accelerate the heated propellant;

Electrostatic methods elements with smaller molar weight are preferred, like hydrogen or hydrazine, but solids are not excluded as well; etc.). Electric arc is mostly used to heat the propellant. Specific impulses of such engines are not very high, compared to other electric propulsion engines (~1000 sec), but are certainly higher than that of chemical engines. The first tests of such engines were performed in mid 1960th on the Soviet Zond-2 Martian probe (the probe has not reached its target, but the thrusters were successfully tested) and the Zond-3 lunar fly-by probe. The engines worked in impulse regime. Later similar devices were used for control on the Meteor and other satellites (helium and argon propellants were used since they are chemically inert and have relatively low ionization potential), as well as on several US satellites. In electrostatic methods, Coulomb forces are used to accelerate ions – these are so-called ion thrusters. Atoms of the working fluid are introduced to the discharge chamber where they get ionized by electrons supplied by an electron gun. The ions drift to the extraction systems containing two grids, positive and negative, and are accelerated between the grids. The negative grid prevents electrons from streaming back to the discharge plasma. Magnetic fields inside the discharge chamber retain electrons increasing their life-time inside the chamber and thus increasing ionizing efficiency. Electrons are also emitted from a separate cathode (neutralizer) outside the vehicle to prevent the vehicle from gaining a negative charge. The elements of the engine are subject to continuous ion bombardment and may corrode, so a suitable propellant should be chosen. Nowadays xenon is mostly used, xenon-based engines have demonstrated intermittent work during years. The values of specific impulse of such devices reach ~ 3000  10 000 sec. The performance may be improved if two pairs of

Electromagnetic methods grids are used, Isp may reach 20 000 sec and higher. A variant of ion thrusters is Field Emission Electric Propulsion, where a very strong electric field extracts ions from the surface of liquid metal propellant and accelerates it. Possessing very high specific impulse (Isp > 10 000 sec), this design provides very low thrust (about 10-6 10-3newtons), so it is applicable generally for attitude control. In another variant (Colloid thrusters), electric field accelerates charged droplets of working liquid. In this group of methods, magnetic field acts to accelerate the working fluid, either through the Lorentz force either through focusing the beam. Quite different methods have been proposed, several designs have been already applied. The operational method of the Hall Effect Thrusters is similar to that of the ion thrusters, but the attractive negative charge is formed by an electron plasma instead of the grid. Ionization of the propellant atoms occurs due to collision with high energy electrons which are trapped by the Hall current; this design significantly increases efficiency of ionization. Such thrusters were first introduced in the USSR and were used on several satellites; nowadays they are introduced worldwide. For instance, the experimental ESA lunar probe SMART-1 was propelled by Hall Effect Thrusters with Isp = 1650 sec, power consumption 1.2 kW. Released at GTO in Sep 2003, it raised its apogee by Oct 2004 so that it passed to a polar lunar orbit and lowerd its aposelene from ~ 55 000 km to ~ 4600 km afterwards. A number of promising methods is being developed which may soon become operational.

Cold gas thrusters Cold gas thrusters is the most simple method of rocket propulsion. A thruster assembly is a vessel containing gas (usually nitrogen) under high pressure, which is connected by a line with a nozzle. Opening the valve leads to gas flow through the nozzle and the reaction force appears. Heaters may be included to provide higher gas temperatures and thus higher efficiency. This design have very low values of specific impulse, Isp ~ 50  100sec, thrust is low as well (tens or hundreds of newton), but due to its simplicity and reliability it is widely used on spacecraft as an attitude control method (primary or supplemental). The mass of the assembly is quite low, mostly it is composed by the masses of the pressurized vessels and the gas. Gas thrusters were used for attitude control on the Luna-3 lunar probe, on the Skylab space station (as supplemental unit; primary attitude control devices were gyros), on the Hipparcos astronomical satellite (also as supplemental unit), and on many others. It is also applied for human individual maneuvering in space, for instance, on the Manned Maneuvering Unit (MMU).

Solar sail Solar sail is the method of propulsion where the radiation pressure of the solar light is used to accelerate or decelerate the spacecraft. A general concept of a spacecraft with a solar sail is a large light-weight aluminum-covered membrane mirror with the spacecraft attached in its center. There are various methods considered to keep the shape of the sail: a light-weight frame structure or a frameless rotational stabilization. Currently films with thickness of several micrometers are available, so the area density of such membranes may be several grams or tens of grams per square meter. The force exerted by solar radiation on a square meter is  being the overall surface reflectance ( = 0  1),  being the angle of incidence, REarthis the radius of the Earth’s heliocentric orbit, R is the distance of the probe From the Sun. So, in the vicinity of the Earth the acceleration of a sail with the area density of 10 g/m2 may achieve ~ 1.5·10-4g. Lighter films would permit to achieve higher accelerations, but, of course, the presence of the payload obliges to increase the area of the sail. Designs with sail dimensions of several kilometers have been analyzed. Changing the inclination of the sail in respect to the direction to the Sun, one would change the angle of incidence and thus, the direction of the reaction force. Of course, the thrust would change as well; it has the maximum value when the light falls onto the surface along its normal. To control the attitude of the spacecraft, additional smaller sails are proposed to act as vanes which would rotate the spacecraft. There is a popular misconception that a solar sail can only accelerate its payload out from the Sun. Of course it is not so: it is also possible to decelerate towards the Sun by pointing the thrust oppositely to the orbital

motion. This would reduce the orbital speed and will bring the sail closer to the Sun. The operational limitation of the solar sail is the distance from the Earth: at the heights >800 km the radiation pressure is smaller than the drag. So solar sails may be used from heights of ~ 1000 km. Creation of extensive ultra-thin films, their compact packing for launch and latter deployment is also complicated from the technical point of view. One attempt to test solar sails in space was made (Cosmos 1, 2005), but it failed due a launch vehicle failure. Some deployment tests have been successfully performed, others have failed due to launch vehicles failures or deployment problems. In general, only small-scale studies are being performed at the present time, no large-scale models have been built. At present, solar sail is regarded as a possible alternative propulsion method for robotic exploration of the Solar system.

End of the Lecture 11

A NERVA solid core design The NERVA engine was designed as the upper stage for the Saturn V(By source)

NERVA stage Vac. thrust ~90 tons Isp ~ 825 sec Mass 34 tons Burn time 20 min Height 43.7 m Diameter 10.6 m (By source) Solid core nuclear engines NERVA & RD-0410 RD-0410 Vac. thrust 3.6 tons Isp ~ 925 sec Mass 2 tons Burn time 1 h Restarts 10 Height 3.7 m Diameter 1.2 m (By source)

Project Orion One of the possible configurations of the spacecraft in the project Orion(By source)

Electrostatic ion thruster DS1 The ion thruster DS1, set on the experimental space probe Deep Space 1. Power: 2.3 kW Thrust: 90 mN Isp: 4000 sec (By source) DS1 at work(By source)

Manned Maneuvering Unit Bruce McCandless in the MMU during a Shuttle’s flight (By source) MMU scheme (By source)

Solar Sail Concepts Frame-stabilized (first) and spin-stabilized solar sails (By source) NanoSail-D experimental solar sail. Lost in launch failure of the Falcon 1 test launch in Aug 2008 (By source)

Mass: 100 kg Blades: 8 Blade length: 15 m Blade area: 600 m2 Material: aluminized reinforced mylar Cosmos 1 solar sail experiment (By source)

Introduction to Astronautics Sissejuhatus kosmonautikasse