Rocket fuel

Rocket fuel is the propellant which is burned to produce thrust in rockets. The term rocket propellant is also used. It shows more unambiguously that not only the fuel proper, but also the oxidizer is meant.

History

Overview

The velocity that a rocket can attain is primarily a function of its mass fraction and its exhaust velocity. This is known as the rocket equation: . The mass fraction is just a way to express how much of the rocket is fuel when it starts accelerating. Typically, a single-stage rocket might be 80% fuel, which is a mass fraction of 5. The exhaust velocity is often reported as specific impulse.

The first stage will usually use high-density (low volume) propellants to reduce the amount of volume exposed to atmospheric drag. 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 propellants

Solid fuels were the first type of propellant to be used in rockets. Gunpowder, obviously, was the original propellant to be used in rocketry, consisting of a mixture of charcoal, sulfur and potassium nitrate (saltpeter). Solid fuels (and really, all rocket fuels) consist of an oxidizer (substance providing oxygen) and a fuel. In the case of gunpowder, the fuel is charcoal and sulfur and the oxidizer is the potassium nitrate. More contemporary recipies employ such compounds as sodium or potassium chlorate and powdered aluminum. (This mixture is sometimes known as "white powder"; not only is it different in appearance than black powder, it has a considerably higher energy density.)

However, white powder has insufficient specific impulse for orbital or near-orbital boosters. 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 aluminum powder (a fuel), held together in a base of PBAN or HTPB (rubber-like fuels). The mixture is formed as a liquid at elevated temperatures, poured into the rocket casing, and cools to form a single grain bonded to that casing.

Solid fueled rockets are much easier to store and handle than liquid fueled rockets, which makes them ideal for military applications. The LGM-30 Minuteman and LG-118A Peacekeeper (MX) missiles are four-stage rockets capable of intercontinental suborbital flights. The first three stages are solid fuelled, and in each case the last stage is a precision maneuverable liquid-fuelled bus used to fine tune the trajectory of the reentry vehicle.

Their simplicity 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. (The Orbital Sciences' Pegasus rocket is a three-stage solid rocket orbital booster.)

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 targetting 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.

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, and the fuel tanks can be built with less material, 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 fuels is the oxidizer. Several practical liquid oxidizers (liquid oxygen, nitrogen tetraoxide) are available which have much better specific impulse than ammonium perchlorate when paired with comparable fuels.

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

The main difficulties with liquid fuels are also with the oxidizers. These are generally difficult to store and handle, either due to extreme toxicity (nitric acids), extreme cold (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 troublesome and highly stressed pressurization systems, plumbing and combustion chambers, which greatly increase the cost of the rocket. Many employ turbopumps which raise the cost still more.

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 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 exaust thermodynamics (making this a tripropellant). The combination delivered 542 seconds (542 lbf·s/lb, 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 trashes 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. Finally, both lithium and fluorine are expensive and rare, enough to actually matter.

The common liquid fuel combinations in use today are:

• LOX and kerosene (RP-1). Used for the lower stages of most Russian and Chinese boosters, and the first stage of the Saturn 5. Very similar to Robert Goddard's first rocket. This combination is widely regarded as the most practical for civilian orbital launches.

• LOX and liquid hydrogen, used in the Space Shuttle, the Centaur upper stage, and the newer Delta IV rocket.

• Nitrogen tetroxide (N₂O₄) and hydrazine (N₂H₄). Used in military, orbital and deep space rockets, because both liquids are storable for long periods at reasonable temperatures and pressures. Hydrazine decomposes energetically to nitrogen and hydrogen, making it a fairly good monopropellant all by itself.
  

  Hybrid propellants

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 fast-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 enough better than hybrids that 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. BYU, 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.

• 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.
  

  Mixture ratio

The usual explanation for fuel-rich mixtures is that fuel-rich
mixtures have lower molecular weight exhaust, which by reducing
increases the ratio ,
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
decreases as fast or faster than .

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 stochiometric of 3.4 to 4), because the energy release
per unit mass drops off quickly as the mixture ratio deviates from
stochiometric. LOX/LH₂ rockets are run very rich (O/F mass
ratio of 4 rather than stochiometic 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-stochiometric mixtures
burn cooler than stochiometric 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.

Nuclear propulsion

Ion propulsion

Fusion propulsion

Photon propulsion

See also