Rocket propellant

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Rocket propellants undergo exothermic chemical reactions which produce hot gas which is used by a rocket for propulsive purposes.

1 Overview
2 Solid propellants
3 Liquid propellants
4 Hybrid propellants
5 Mixture ratio
6 Propellant density
7 See also
8 External links

[edit] Overview

Rockets create thrust by expelling mass backwards in a high speed jet (see Newton's Third Law ). Chemical rockets, the subject of this article, create thrust by reacting propellants into very hot gas, which then expands and accelerates within a nozzle out the back. The thrust produced is the mass flow rate of the propellants multiplied by their exhaust velocity (relative to the rocket), as specified by Newton's third law of motion. It is the equal and opposite reaction that moves the rocket, and not any interaction of the exhaust stream with air around the rocket (but see base bleed). Equivalently, one can think of a rocket being accelerated upwards by the pressure of the combusting gases in the combustion chamber and nozzle. Rockets perform better in space, because they do not need to overcome air resistance and atmospheric pressure on the outside of the nozzle.

The maximum velocity that a rocket can attain in the absence of any external forces is primarily a function of its mass ratio and its exhaust velocity. The relationship is described by the rocket equation: $V f = V e ln(M 0 / M f)$ . The mass ratio is just a way to express what proportion of the rocket is fuel when it starts accelerating. Typically, a single-stage rocket might have a mass fraction of 90% propellant, which is a mass ratio of 1/(1-0.9) = 10. The exhaust velocity is often reported as specific impulse.

The first stage will usually use high-density (low volume) propellants to reduce the area exposed to atmospheric drag and because of the lighter tankage and higher thrust/weight ratios. Thus, the Apollo-Saturn V first stage used kerosene-liquid oxygen rather than the liquid hydrogen-liquid oxygen used on the upper stages (hydrogen is highly energetic per kilogram, but not per cubic metre). Similarly, the Space Shuttle uses high-thrust, high-density SRBs for its lift-off with the liquid hydrogen-liquid oxygen SSMEs used partly for lift-off but primarily for orbital insertion.

There are three main types of propellants: solid, liquid, and hybrid.

[edit] Solid propellants

Main article: Solid-fuel rocket

The earliest rockets were created hundreds of years ago by the Chinese, and were used primarily for fireworks displays and as weapons. They were fueled with black powder, a type of gunpowder consisting of a mixture of charcoal, sulfur and potassium nitrate (saltpeter). Rocket propellant technology did not advance until the end of the 19th century, by which time smokeless powder had been developed, originally for use in firearms and artillery pieces. Smokeless powders and related compounds have seen use as double-base propellants.

Solid propellants (and almost all rocket propellants) consist of an oxidizer and a fuel. In the case of gunpowder, the fuel is charcoal, the catalyst is sulfur (Note: sulfur is not a true catalyst in gunpowder as it is consumed to a great extent into a variety of reaction products such as K₂S. The sulfur however acts mainly as a sensitizer lowering threshold of ignition) and the oxidizer is the potassium nitrate. During the 1950s and 60s researchers in the United States developed what is now the standard high-energy solid rocket fuel. The mixture is primarily ammonium perchlorate powder (an oxidizer), combined with fine aluminium powder (a fuel), held together in a base of PBAN or HTPB (rubber-like fuels). The mixture is formed as a liquid, and then cast into the correct shape and cured into a rubbery solid. Solid fueled rockets are much easier to store and handle than liquid fueled rockets, which makes them ideal for military applications. In the 1970s and 1980s the U.S. switched entirely to solid-fuelled ICBMs: the LGM-30 Minuteman and LG-118A Peacekeeper (MX). In the 1980s and 1990s, the USSR/Russia also deployed solid-fuelled ICBMs (RT-23, RT-2PM, and RT-2UTTH), but retains two liquid-fuelled ICBMs (R-36 and UR-100N). All solid-fuelled ICBMs on both sides have three initial solid stages and a precision maneuverable liquid-fuelled bus used to fine tune the trajectory of the reentry vehicle.

Their simplicity also makes solid rockets a good choice whenever large amounts of thrust are needed and cost is an issue. The Space Shuttle and many other orbital launch vehicles use solid fuelled rockets in their first stages (solid rocket boosters) for this reason.

However, solid rockets have lower specific impulse than liquid fueled rockets. It is also difficult to build a large mass ratio solid rocket because almost the entire rocket is the combustion chamber, and must be built to withstand the high combustion pressures. If a solid rocket is used to go all the way to orbit, the payload fraction is very small. (For example, the Orbital Sciences Pegasus rocket is an air-launched three-stage solid rocket orbital booster. Launch mass is 23,130 kg, low earth orbit payload is 443 kg, for a payload fraction of 1.9%. Compare to a Delta IV Medium, 249,500 kg, payload 8600 kg, payload fraction 3.4% without air-launch assistance.)

Solid rockets are difficult to throttle or shut down before they run out of fuel. Essentially, the burning grain must be vented to lower the chamber pressure. Venting generally involves destroying the rocket, and is usually only done by a Range Safety Officer if the rocket goes awry. The third stages of the Minuteman and MX rockets have precision shutdown ports which, when opened, reduce the chamber pressure so abruptly that the interior flame is blown out. This allows a more precise trajectory which improves targeting accuracy.

Finally, casting very large single-grain rocket motors has proved to be a very tricky business. Defects in the grain can cause explosions during the burn, and these explosions can increase the burning propellant surface enough to cause a runaway pressure increase, until the case fails.

[edit] Liquid propellants

Main article: Liquid rocket propellants

Liquid fueled rockets have better specific impulse than solid rockets and are capable of being throttled, shut down, and restarted. Only the combustion chamber of a liquid fueled rocket needs to withstand combustion pressures and temperatures. On vehicles employing turbopumps, the fuel tanks carry very much less pressure and thus can be built far more lightly, permitting a larger mass fraction. For these reasons, most orbital launch vehicles and all first- and second-generation ICBMs use liquid fuels for most of their velocity gain.

The primary performance advantage of liquid propellants is the oxidizer. Several practical liquid oxidizers (liquid oxygen, nitrogen tetroxide, and hydrogen peroxide) are available which have much better specific impulse than ammonium perchlorate when paired with comparable fuels.

Most liquid propellants are also cheaper than solid propellants. For orbital launchers, the cost savings do not, and historically have not mattered; the cost of fuel is a very small portion of the overall cost of the rocket, even in the case of solid fuel.

The main difficulties with liquid propellants are also with the oxidizers. These are generally at least moderately difficult to store and handle, either due to extreme toxicity (nitric acids), moderately cryogenic (liquid oxygen), or both (liquid fluorine is a perennial favorite of wild-eyed enthusiasts). Several exotic oxidizers have been proposed: liquid ozone (O₃), ClF₃, and ClF₅, all of which are unstable, energetic, and toxic.

Liquid fuelled rockets also require potentially troublesome valves and seals and thermally stressed combustion chambers, which increase the cost of the rocket. Many employ specially designed turbopumps which raise the cost enormously due to difficult fluid flow patterns that exist within the casings.

Though all the early rocket theorists proposed liquid hydrogen and liquid oxygen as propellants, the first liquid-fuelled rocket, launched by Robert Goddard on March 16, 1926, used gasoline and liquid oxygen. Liquid hydrogen was first used by the engines designed by Pratt and Whitney for the Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950s. In the mid-1960s, the Centaur and Saturn upper stages were both using liquid hydrogen and liquid oxygen.

The highest specific impulse chemistry ever test-fired in a rocket engine was lithium and fluorine, with hydrogen added to improve the exhaust thermodynamics (making this a tripropellant). The combination delivered 542 seconds (5.32 kN·s/kg, 5320 m/s) specific impulse in a vacuum. The impracticality of this chemistry highlights why exotic propellants are not actually used: to make all three components liquids, the hydrogen must be kept below -252 °C (just 21 K) and the lithium must be kept above 180 °C (453 K). Lithium and fluorine are both extremely corrosive, lithium ignites on contact with air, fluorine ignites on contact with most fuels, and hydrogen, while not hypergolic, is an explosive hazard. Fluorine and the hydrogen fluoride (HF) in the exhaust are very toxic, which damages the environment, makes work around the launch pad difficult, and makes getting a launch license that much more difficult. The rocket exhaust is also ionized, which would interfere with radio communication with the rocket.

The common liquid propellant combinations in use today:

LOX and kerosene (RP-1). Used for the lower stages of most Russian and Chinese boosters, and the first stage of the Saturn V. Very similar to Robert Goddard's first rocket. This combination is widely regarded as the most practical for civilian orbital launchers.

LOX and liquid hydrogen, used in the Space Shuttle, the Centaur upper stage, the newer Delta IV rocket, and most stages of the European Ariane rockets.

Nitrogen tetroxide (N₂O₄) and hydrazine (N₂H₄), MMH, or UDMH. Used in military, orbital and deep space rockets, because both liquids are storable for long periods at reasonable temperatures and pressures. This combination is hypergolic, making for attractively simple ignition sequences. The major inconvenience is that these propellants are highly toxic, hence they require careful handling. Hydrazine decomposes energetically to nitrogen and hydrogen, making it a fairly good monopropellant all by itself.

[edit] Hybrid propellants

A hybrid rocket usually has a solid fuel and a liquid or gas oxidizer. The fluid oxidizer can make it possible to throttle and restart the motor just like a liquid fuelled rocket. Hybrid rockets are also cleaner than solid rockets because practical high-performance solid-phase oxidizers all contain chlorine, versus the more benign liquid oxygen or nitrous oxide used in hybrids. Because just one propellant is a fluid, hybrids are simpler than liquid rockets.

Hybrid motors suffer two major drawbacks. The first, shared with solid rocket motors, is that the casing around the fuel grain must be built to withstand full combustion pressure and often extreme temperatures as well. Modern composite structures handle this problem well.

The primary remaining difficulty with hybrids is with mixing the propellants before burning. In solid propellants, the oxidizer and fuel are mixed in a factory in carefully controlled conditions (and even then it is tricky). Liquid propellants are generally mixed by the injector at the top of the combustion chamber, which directs many small swift-moving streams of fuel and oxidizer into one another. Liquid fuelled rocket injector design has been studied at great length and still resists reliable performance prediction. In a hybrid motor, the mixing happens at the surface of the melting or evaporating surface of the fuel. The mixing is not a well-controlled process and generally quite a lot of propellant is left unburned, which limits the efficiency and thus the exhaust velocity of the motor.

There has been much less development of hybrid motors than solid and liquid motors. For military use, ease of handling and maintenance have driven the use of solid rockets. For orbital work, liquid fuels are more efficient than hybrids and most development has concentrated there. There has recently been an increase in hybrid motor development for nonmilitary suborbital work:

The Reaction Research Society (RRS), although known primarily for their work with liquid rocket propulsion, has a long history of research and development with hybrid rocket propulsion.

Several universities have recently experimented with hybrid rockets. Brigham Young University, the University of Utah and Utah State University launched a student-designed rocket called Unity IV in 1995 which burned the solid fuel hydroxy-terminated polybutadiene (HTPB) with an oxidizer of gaseous oxygen, and in 2003 launched a larger version which burned HTPB with nitrous oxide.

Portland State University also launched several hybrid rockets in the early 2000's.

The Rochester Institute of Technology is currently creating a HTPB hybrid rocket to launch small payloads into space and to several near Earth objects. Its first launch is scheduled for Summer 2007. http://meteor.rit.edu

Scaled Composites SpaceShipOne, the first private manned spacecraft, is powered by a hybrid rocket burning HTPB with nitrous oxide. The hybrid rocket engine was manufactured by SpaceDev. SpaceDev partially based its motors on experimental data collected from the testing of AMROC's (American Rocket Company) motors at NASA's Stennis Space Center's E1 test stand. Motors ranging from as small as 1000 lbf (4.4 kN) to as large as 250,000 lbf (1.1 MN) thrust were successfully tested. SpaceDev purchased AMROCs assets after the company was shut down for lack of funding.

[edit] Mixture ratio

The theoretical exhaust velocity of a given propellant chemistry is a function of the energy released per unit of propellant mass (specific energy). Unburned fuel or oxidizer drags down the specific energy. Surprisingly, most rockets run fuel-rich.

The usual explanation for fuel-rich mixtures is that fuel-rich mixtures have lower molecular weight exhaust, which by reducing $M$ supposedly increases the ratio $\frac{\sqrt{T_c}}{M}$ which is approximately equal to the theoretical exhaust velocity. This explanation, though found in some textbooks, is wrong. Fuel-rich mixtures actually have lower theoretical exhaust velocities, because $\sqrt{T_c}$ decreases as fast or faster than $M$ .

The nozzle of the rocket converts the thermal energy of the propellants into directed kinetic energy. This conversion happens in a short time, on the order of one millisecond. During the conversion, energy must transfer very quickly from the rotational and vibrational states of the exhaust molecules into translation. Molecules with fewer atoms (like CO and H₂) store less energy in vibration and rotation than molecules with more atoms (like CO₂ and H₂O). These smaller molecules transfer more of their rotational and vibrational energy to translation energy than larger molecules, and the resulting improvement in nozzle efficiency is large enough that real rocket engines improve their actual exhaust velocity by running rich mixtures with somewhat lower theoretical exhaust velocities.

The effect of exhaust molecular weight on nozzle efficiency is most important for nozzles operating near sea level. High expansion rockets operating in a vacuum see a much smaller effect, and so are run less rich. The Saturn-II stage (a LOX/LH₂ rocket) varied its mixture ratio during flight to optimize performance.

LOX/hydrocarbon rockets are run only somewhat rich (O/F mass ratio of 3 rather than stoichiometric of 3.4 to 4), because the energy release per unit mass drops off quickly as the mixture ratio deviates from stoichiometric. LOX/LH₂ rockets are run very rich (O/F mass ratio of 4 rather than stoichiometric 8) because hydrogen is so light that the energy release per unit mass of propellant drops very slowly with extra hydrogen. In fact, LOX/LH₂ rockets are generally limited in how rich they run by the performance penalty of the mass of the extra hydrogen tankage, rather than the mass of the hydrogen itself.

Another reason for running rich is that off-stoichiometric mixtures burn cooler than stoichiometric mixtures, which makes engine cooling easier. And as most engines are made of metal or carbon, hot oxidizer-rich exhaust is extremely corrosive, where fuel-rich exhaust is less so. American engines have all been fuel-rich. Some Soviet engines have been oxidizer-rich.

Additionally, there is a difference between mixture ratios for optimum specific impulse (I_sp) and optimum thrust. During launch, shortly after takeoff, high thrust is at a premium. This can be achieved at some temporary reduction of Isp by increasing the oxidiser ratio initially, and then transitioning to more fuel-rich mixtures. Since engine size is typically scaled for takeoff thrust this permits reduction of the weight of rocket engine, pipes and pumps and the extra propellant use can be more than compensated by increases of acceleration towards the end of the burn by having a reduced dry mass.

[edit] Propellant density

Although liquid hydrogen gives a high Isp, its low density is a significant disadvantage: hydrogen occupies about 7x more volume per kilogram than dense fuels such as kerosene. This not only penalises the tankage, but also the pipes and fuel pumps leading from the tank, which need to be 7x bigger and heavier. (The oxidiser side of the engine and tankage is of course unaffected.) This makes the vehicle's dry mass much higher, so the use of liquid hydrogen is not such a big win as might be expected. Indeed, some dense hydrocarbon/LOX propellant combinations have higher performance when the dry mass penalties are included.

Due to lower Isp, dense propellant launch vehicles have a higher takeoff mass, but this does not mean a proportionately high cost; on the contrary, the vehicle may well end up cheaper. Liquid hydrogen is quite an expensive fuel to produce and store, and causes many practical difficulties with design and manufacture of the vehicle.

Because of the higher overall weight, a dense-fuelled launch vehicle necessarily requires higher takeoff thrust, but it carries this thrust capability all the way to orbit. This, in combination with the better thrust/weight ratios, means that dense-fuelled vehicles reach orbit earlier, thereby minimizing losses due to gravity drag. Thus, the effective delta-v requirement for these vehicles are reduced.

However, liquid hydrogen does give clear advantages when the overall mass needs to be minimised; for example the Saturn V vehicle used it on the upper stages; this reduced weight meant that the dense-fuelled first stage could be made proportionately smaller, saving quite a bit of money.